Propulsor trim prediction for aircraft

ABSTRACT

A control circuitry includes a first filter configured to filter a gravity compensated longitudinal acceleration of an aircraft to generate a filtered gravity compensated longitudinal acceleration. The propulsor trim control circuitry also includes a second filter configured to generate a filtered speed of the aircraft based on a speed of the aircraft. The propulsor trim control circuitry includes intermediary circuitry configured to generate a filtered longitudinal control effector error based on the filtered gravity compensated longitudinal acceleration and the speed. The propulsor trim control circuitry also includes a third filter configured to generate a filtered longitudinal thrust effector command value based on a longitudinal thrust effector command value. The propulsor trim control circuitry further includes output circuitry configured to generate a predicted longitudinal thrust effector trim value for a target horizontal state based on the filtered longitudinal control effector error and the filtered longitudinal thrust effector command value.

CROSS-REFERENCE

This application claims priority from and the benefit of U.S.Provisional Patent Application No. 62/662,674, filed Apr. 25, 2018 andentitled “PITCH AND THRUST CONTROL FOR AIRCRAFT”, the contents of whichare expressly incorporated herein by reference in their entirety.

This application is related to U.S. patent application Ser. No.15/986,581, filed on May 22, 2018 and entitled “PITCH AND THRUST CONTROLFOR COMPOUND AIRCRAFT”, the contents of which are expressly incorporatedherein by reference in their entirety.

This application is related to U.S. patent application Ser. No.15/986,606, filed on May 22, 2018 and entitled “PITCH AND THRUST CONTROLFOR TILT-ROTOR AIRCRAFT”, the contents of which are expresslyincorporated herein by reference in their entirety.

This application is related to U.S. patent application Ser. No.15/986,634, issued as U.S. Pat. No. 10,747,235, filed on May 22, 2018and entitled “PITCH TRIM PREDICTION FOR AIRCRAFT”, the contents of whichare expressly incorporated herein by reference in their entirety.

FIELD OF THE DISCLOSURE

The present disclosure is generally related to propulsor trim predictionfor aircraft.

BACKGROUND

Conventional rotorcraft (e.g., helicopters) require substantial pilotskill and workload for operation. As such, rotorcraft operators expendsignificant resources on pilot training and proficiency. High-speedVertical Take-Off and Landing (VTOL) aircraft, such as compound aircraft(e.g., compound helicopters), tilt-rotor aircraft, and jump jets, areamong the most complicated aircraft to pilot and require extensivetraining.

Such high-speed VTOL aircraft commonly require the pilot tosimultaneously use five control inputs to control (indirectly control)two output states (e.g., airspeed and climb rate). For example, atilt-rotor aircraft includes a first input to control pitch moment, asecond input to control yaw moment, a third input to control rollmoment, a fourth input to control proprotor thrust, and a fifth input tocontrol proprotor shaft orientation. The pilot may have to coordinateall of these controls simultaneously to operate the aircraft.Additionally, such compound aircraft may include multiple trim solutionsfor specific states that are not intuitive, i.e., multiple solutions mayexist to balance the moments and the thrust to achieve a desired state.For example, a particular trim solution may increase or minimize fuelburn, while another trim solution may reduce or minimize noise, and yetanother trim solution may increase or maximize agility. Thus, withconventional controls the pilot cannot easily reconfigure the aircraftas the mission demands.

SUMMARY

In a particular implementation, a control circuitry includes a propulsortrim prediction circuitry and output circuitry. The propulsor trimprediction circuitry is configured to generate a predicted propulsorcollective blade pitch trim value for an aircraft based on an aircraftvelocity and a pitch attitude deviation from a reference. The outputcircuitry is configured to output a propulsor collective blade pitchangle command based on the predicted propulsor collective blade pitchtrim value. The propulsor collective blade pitch angle command isconfigured to cause an adjustment in a collective blade pitch angle of apropulsor of the aircraft.

In another particular implementation, a control circuitry includes apropulsor trim prediction circuitry and output circuitry. The propulsortrim prediction circuitry is configured to generate a predictedproprotor nacelle trim value based on an aircraft velocity and a pitchattitude deviation from a reference. The output circuitry is configuredto output a proprotor nacelle command based on the predicted proprotornacelle trim value. The proprotor nacelle command is configured to causean adjustment in a nacelle angle of a proprotor of an aircraft.

In another particular implementation, a control circuitry includes apitch attitude trim prediction circuitry and output circuitry. The pitchattitude trim prediction circuitry is configured to generate a predictedpitch attitude trim value for an aircraft based on an aircraft velocityand a pitch attitude of the aircraft. The output circuitry is configuredto output an aircraft pitch attitude trim command based on the predictedpitch attitude trim value and a pilot input signal from a pitch controlinceptor. The aircraft pitch attitude trim command is configured tocause an adjustment in a pitch angle of the aircraft.

In a particular implementation, a propulsor trim control circuitryincludes a first filter configured to filter a gravity compensatedlongitudinal acceleration of an aircraft to generate a filtered gravitycompensated longitudinal acceleration. The propulsor trim controlcircuitry also includes a second filter configured to generate afiltered speed of the aircraft based on a speed of the aircraft. Thepropulsor trim control circuitry includes intermediary circuitryconfigured to generate a filtered longitudinal control effector errorbased on the filtered gravity compensated longitudinal acceleration andthe filtered speed. The propulsor trim control circuitry also includes athird filter configured to generate a filtered longitudinal thrusteffector command value based on a longitudinal thrust effector commandvalue. The propulsor trim control circuitry further includes outputcircuitry configured to generate a predicted longitudinal thrusteffector trim value for a target horizontal state based on the filteredlongitudinal control effector error and the filtered longitudinal thrusteffector command value. The predicted longitudinal thrust effector trimvalue is configured to cause a longitudinal thrust effector of theaircraft to be adjusted.

In a particular implementation, a pitch trim prediction circuitryincludes a first filter configured to generate a filtered velocity basedon a component of a vertical velocity of an aircraft. The pitch trimprediction circuitry also includes a second filter configured togenerate a filtered pitch attitude based on a measured pitch attitude ofthe aircraft. The pitch trim prediction circuitry further includesoutput circuitry configured to generate a predicted pitch attitude trimvalue for a target vertical state based on a horizontal velocity of theaircraft, the filtered velocity, and the filtered pitch attitude. Thepredicted pitch attitude trim value is configured to cause a flightcontrol effector to be adjusted.

In a particular implementation, a method for controlling an aircraftincludes generating a predicted propulsor collective blade pitch trimvalue for a target state of the aircraft based on an aircraft velocityand a pitch attitude deviation from a reference. The method furtherincludes adjusting a propulsor collective blade pitch angle of apropulsor of the aircraft based on the predicted propulsor collectiveblade pitch trim value.

In a particular implementation, a method for controlling an aircraftincludes generating a predicted proprotor nacelle trim value for atarget state of the aircraft based on an aircraft velocity and a pitchattitude deviation from a reference. The method further includesadjusting a nacelle angle of a proprotor of the aircraft based on thepredicted proprotor nacelle trim value.

In another particular implementation, a method for controlling anaircraft includes generating a predicted pitch attitude trim value for atarget state of an aircraft based on an aircraft velocity and a pitchattitude of the aircraft. The method also includes adjusting an aircraftpitch attitude command based on the predicted pitch attitude trim valueand a pilot input signal from a pitch control inceptor.

In another particular implementation, a method of controlling anaircraft includes receiving a vertical velocity of the aircraft and ahorizontal velocity of the aircraft. The method also includes filteringa component of the vertical velocity of the aircraft to generate afiltered vertical velocity. The method includes filtering a measuredpitch attitude of the aircraft to generate a filtered pitch attitude.The method also includes generating a predicted pitch attitude trimvalue for a target vertical state, the predicted pitch attitude trimvalue generated based on the horizontal velocity, the filtered verticalvelocity, and the filtered pitch attitude. The method further includesadjusting a flight control effector based on the predicted pitchattitude trim value.

In another particular implementation, a method of controlling anaircraft includes filtering a gravity compensated longitudinalacceleration of the aircraft to generate a filtered gravity compensatedlongitudinal acceleration. The method also includes filtering a speed ofthe aircraft to generate a filtered speed of the aircraft. The methodincludes generating a filtered longitudinal control effector error basedon the filtered gravity compensated longitudinal acceleration and thefiltered speed. The method also includes filtering a longitudinal thrusteffector command value to generate a filtered longitudinal thrusteffector command value. The method includes generating a predictedlongitudinal thrust effector trim value for a target horizontal statebased on the filtered longitudinal control effector error and thefiltered longitudinal thrust effector command value. The method furtherincludes adjusting a longitudinal thrust effector of the aircraft basedon the predicted longitudinal thrust effector trim value.

By using pitch trim prediction, propulsor trim prediction, or both,flight control surfaces and thrust effectors of the aircraft can becontrolled with less pilot input and produce more optimized or enhancedflight operations. Accordingly, pilot demand and workload are decreased,which may lead to reduced training, reduced pilot errors, and increasedpilot satisfaction. Additionally, by using pitch trim prediction,propulsor trim prediction, or both, efficiency and/or operationalcapabilities of the aircraft are increased.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a block diagram that illustrates an example of an aircraftincluding a control system;

FIGS. 2A, 2B, 2C, and 2D are each a diagram that illustrates controls ofa compound helicopter;

FIGS. 3A, 3B, 3C, and 3D are each a diagram that illustrates controls ofa tiltrotor aircraft;

FIG. 4 is a circuit diagram that illustrates an example of the controlcircuitry of FIG. 1;

FIG. 5 is a circuit diagram that illustrates another example of thecontrol circuitry of FIG. 1;

FIG. 6 is a circuit diagram that illustrates another example of thecontrol circuitry of FIG. 1;

FIG. 7 is a circuit diagram that illustrates an example of a pitch trimprediction circuitry;

FIG. 8 is a circuit diagram that illustrates another example of a pitchtrim prediction circuitry;

FIG. 9 is a circuit diagram that illustrates an example of a pitchcommand model;

FIG. 10 is a circuit diagram that illustrates an example of a speedselect circuitry;

FIG. 11 is a circuit diagram that illustrates an example of anacceleration command circuitry;

FIG. 12 is a circuit diagram that illustrates an example of anacceleration controller;

FIG. 13 is a circuit diagram that illustrates an example of a propulsortrim prediction circuitry;

FIG. 14 is a circuit diagram that illustrates an example of a propulsorlimiting circuitry;

FIG. 15 is a flow chart of an example of a method for controlling anaircraft;

FIG. 16 is a flow chart of another example of a method for controllingan aircraft;

FIG. 17 is a flow chart of an example of a method for controlling anaircraft;

FIG. 18 is a flow chart of an example of a method for controlling anaircraft;

FIG. 19 is a flow chart of an example of a method of controlling anaircraft;

FIG. 20 is a flow chart of an example of a method of controlling anaircraft;

FIG. 21 is a flow chart of an example of a method of operating a controlcircuitry manufacturing system; and

FIG. 22 is a block diagram that illustrates an example of a vehicleincluding a control circuitry.

DETAILED DESCRIPTION

Implementations disclosed herein include control circuitry for compoundaircraft having pitch trim prediction circuitry, propulsor trimprediction circuitry, or both. Such control circuitry enables moreefficient control of the compound aircraft and improved performance withfewer pilot inputs (or autopilot inputs). The pitch trim predictioncircuitry provides for automatic aircraft (e.g., fuselage) pitch controland automatic main rotor collective control. With the pitch trimprediction circuitry, pilots of compound aircraft fly like a helicopterin a low-speed regime (with a gravity amplifier) and fly like anairplane in a high-speed regime. Additionally, in the high speed regime,the aircraft pitch trim is automatically controlled (e.g., independentof additional pilot control inputs) such that vertical acceleration ismaintained at zero and rate of climb is maintained at a desired value,where the desired value of rate of climb/descent can be zero ornon-zero. Aft cyclic control pressure (e.g., moving the cyclic inceptorrearwards) in high-speed flight commands pitch rate and climb rate withautomatic main rotor and propeller pitch control to hold airspeedconstant and maximize the climb rate. Similarly, forward pressure on aninceptor (e.g., an acceleration thumbwheel) in high-speed flightcommands an increase in airspeed through an increase in longitudinalthrust (e.g., by an increase in propeller collective blade pitch) whilemain rotor collective and aircraft pitch attitude are automaticallycontrolled to maintain a desired rate of climb or descent, where thedesired rate of climb or descent can be zero or non-zero.

The propulsor trim prediction circuitry (e.g., longitudinal thrust trimprediction circuitry) uses commanded propulsor values (e.g., propellercollective blade pitch), inertial sensor data, and dynamic inversioncontrol to calculate propulsor variables for zero acceleration (rate ofchange of speed or v-dot). A propulsor variable can be adjusted to acommand acceleration based on pilot inceptor inputs, thereby providing aflexible inner loop controller that can be coupled to any pilot inceptorand tailored based upon different flight regimes and mission tasks. Thecontrol circuitry overcomes non-linear characteristics introduced bypropeller torque limits, propeller pitch angle limits, propeller pitchactuator rate limits, propeller stall, and propeller reverse torque andovercomes non-linear propeller response near zero torque.

The control circuitry enables increased airspeed-select performance ascompared to conventional solutions, which exhibit overshoots (e.g., fromintegrator wind-up caused by using integral feedback). For example, whena pilot changes states or selects a new value for a current state,conventional control circuitry causes the aircraft to overshoot thedesired value and then undershoot the desired value before settling onthe desired value. During this time, pilots often try to correct theover or undershoots, which further exacerbates the delay until theaircraft achieves and maintains the desired value. Additionally, thecontrol circuitry is also more robust to disturbances and changes indrag caused by external loads, external stores, landing gear deployment,etc.

As compared to conventional control circuitry for compound aircraft, thepilot commands two output states directly, rather than three or moreintermediate states that affect the two output states relevant to flightpath control. With the control circuitry, the pilot is not required tocontrol a fifth control input (e.g., the collective stick) in high-speedflight. This reduces the control strategy to include four control inputsto control 2 (or 4) aircraft states. Similarly, in order to preservehelicopter-like handling qualities in low-speed flight, the pilot is notrequired to control a particular inceptor (e.g., the accelerationthumbwheel) and commands longitudinal acceleration instead throughlongitudinal cyclic stick inputs as in a conventional helicopter. Inorder to retain complete aircraft capability, in some implementationsthe fifth control input is available to the pilot in both modes.However, simultaneously working all five control inputs is not requiredto achieve a desired flight trajectory. For instance, additional inputsare provided to the pilot to command level-body acceleration, level-bodyclimb, and constant speed trim pitch attitude adjustments for specificinstances where this capability is desired. In these instances, thecontrol circuitry reduces the workload associated with coordinatingcontrol inputs to achieve the desired response. Furthermore, inconventional compound aircraft, an airspeed hold is a separate mode thatthe pilot engages when desired, and the airspeed hold mode is not partof the primary flight control system (i.e., is not automatically engagedby the control circuitry based on one or more conditions). Thus, byusing control circuitry with pitch trim prediction, propulsor trimprediction, or both, pilot workload is reduced and the aircraft hasincreased efficiency and improved performance and/or capabilities.

FIG. 1 illustrates a block diagram of an example of an aircraft 100 thatincludes a propulsion system 102 and a control system 104. The controlsystem 104 is configured to control the propulsion system 102 andcontrol surfaces 128 of the aircraft 100. The control system 104receives inputs (e.g., pilot or autopilot inputs), and the controlsystem 104 controls (e.g., directly controls) output states of theaircraft 100 based on the inputs. The aircraft 100 may include orcorrespond to a compound aircraft, such as a compound helicopter, avertical take-off and landing (VTOL) aircraft, a tilt-rotor aircraft,etc. The aircraft 100 may be manned or autonomous.

The propulsion system 102 includes a first propulsor 112, a secondpropulsor 114, and one or more propulsor actuators 116. Although twopropulsors 112, 114 are illustrated in FIG. 1, in other implementations,the propulsion system 102 includes one or more additional propulsors.The propulsors 112, 114 of the propulsion system 102 may includedifferent types of propulsors. For example, in a first implementation,the first propulsor 112 includes or corresponds to a main rotor (e.g., avertical propulsor) of the aircraft 100, and the second propulsor 114corresponds to a propeller (e.g., a longitudinal thruster or propulsor)of the aircraft 100. Alternatively, the second propulsor 114 may includeor correspond to other type of longitudinal thrusters, such as a ductedfan, a contra-rotating fan, a turbojet engine, a turbofan engine, arocket, etc. In other implementations, the first and second propulsors112, 114 are the same type of propulsor. For example, the aircraft 100is a tilt-rotor aircraft, and the first and second propulsors 112, 114correspond to proprotors of the aircraft 100.

The one or more propulsor actuators 116 are coupled to one or more ofthe propulsors 112, 114 and are configured to adjust one or morepropulsors 112, 114 of the propulsion system 102. For example, a firstpropulsor actuator is configured to adjust a collective blade pitchangle of propeller blades of the first propulsor 112. As anotherexample, the first propulsor actuator is configured to adjust a nacelleangle (or a rate of change of the nacelle angle) of the first propulsor112.

The control system 104 includes at least one inceptor device 122, a fullauthority digital engine control (FADEC) 124, a flight control computer(FCC) 126, multiple control surfaces 128, a control circuitry 130, andsensors 132. The inceptor device 122 is configured to control variouscomponents of the aircraft 100. The inceptor device 122 include one ormore inceptors 142, 144 each configured to control a particularcomponent or components of the aircraft 100. As illustrated in FIG. 1,the inceptor device 122 include a first inceptor 142 and a secondinceptor 144. As an illustrative example, the inceptor device 122includes or corresponds to a cyclic stick which is moveable in multipleaxes (e.g., first and second inceptors, one for each axis) and includesa thumbwheel (e.g., a third inceptor). Other inceptor devices includefoot pedals and a collective stick (e.g., fourth and fifth inceptors).

The FADEC 124 is configured to control the propulsion system 102responsive to signals from the inceptor device 122, the FCC 126, or acombination thereof. For example, the signals from the inceptor device122 may be routed through the FCC 126 and processed by the FCC 126. Inother implementations, the control system 104 does not include the FADEC124. The FCC 126 is configured to control operation of the aircraft 100.The FCC is configured to receive signals from one or more components ofthe aircraft 100, process the signals, and outputs commands to one ormore components of the aircraft 100.

The multiple control surfaces 128 are configured to adjust the aircraft100 in flight responsive to signals from the FCC 126. The multiplecontrol surfaces 128 may include or correspond to elevators, ailerons(e.g., flaperons), rudder, flaps, etc. Such control surfaces, generallyare using in high speed flight (e.g., airplane mode) and have lessimpact or effect in low speed flight (e.g., helicopter mode).

The control circuitry 130 is configured to predict an aircraft pitchattitude trim value (a predicted pitch attitude trim value) and apropulsor trim value (a predicted propulsor trim value). The predictedpitch attitude trim value represents an estimated pitch attitude of theaircraft 100 used to trim the aircraft 100 to a particular state (e.g.,an airspeed hold, an altitude hold, a vertical velocity hold, anacceleration hold, etc.). For example, the estimated pitch attitudevalue can be used to determine a setting for a particular controlsurface 128 that will cause the particular control surface 128 to adjustthe pitch attitude of the aircraft 100 to trim the aircraft 100 to aparticular state. The predicted propulsor trim value represents anestimated propulsor value (e.g., an estimated value of a collectiveblade pitch, a nacelle angle, a nacelle angular rate, a nozzle size, anozzle direction, a fuel flow rate, etc.) of one or more of thepropulsors 112, 114 used to trim the aircraft 100 to a particular state.For example, the estimated propulsor value is configured to cause thepropulsor to generate a particular magnitude and/or direction of thrustto trim the aircraft 100 to a particular state. The control circuitry130 may be included in the FCC 126, included in the FADEC 124, or may beseparate from the FADEC 124 and/or the FCC 126.

The sensors 132 are configured to generate sensor data regarding theaircraft 100. For example, the sensors 132 include one or more of apitot static tube, an accelerometer, a gyroscope, etc. The sensorsoutputs sensor data indicating velocity, acceleration, attitudealtitude, etc. The sensors 132 output the sensor data to the FCC 126.

During operation of a compound helicopter (a particular aircraft 100) ina first flight regime (e.g., at a speed below a speed threshold), thepilot (or autopilot) controls the aircraft 100 similar to a conventionalhelicopter. For example, the pilot uses the collective and the inceptordevice 122 (e.g., a cyclic) to fly the aircraft 100. To illustrate, thepilot can use the collective to increase vertical speed by commanding acollective blade pitch angle increase in the first propulsor 112 (e.g.,a main rotor), and the pilot can use the inceptor device 122 (e.g., thecyclic) to increase longitudinal speed by commanding an increase inforward tilt of the first propulsor 112 (e.g., the main rotor) forward.

The aircraft 100 changes flight regimes in response to an increase inspeed above the speed threshold. In the second regime, the pilotcontrols the aircraft 100 similar to a conventional airplane. Forexample, the pilot controls the aircraft 100 by moving the inceptordevice 122 (which controls the control surfaces 128) and by controllingthrust of the second propulsor 114. To illustrate, the pilot controlsthe second inceptor 144 (e.g., a thumbwheel) of the inceptor device 122to increase longitudinal thrust of the second propulsor 114. When thesecond propulsor 114 is a propeller, the propulsor actuator 116 receivesa command to adjust the collective blade pitch of the propeller, whichincreases thrust of the second propulsor 114. As compared toconventional control systems, the aircraft 100 is controlled independentof the collective input device, and the pilot does not have to manage afifth task (e.g., the collective in high speed flight). The pilot canchoose to manage the fifth control input to directly control collectiveblade pitch of the first propulsor 112.

In other implementations, such as where a different type of secondpropulsor 114 is used in conjunction with the main rotor (the firstpropulsor 112), other aspects of the second propulsor 114 may beadjusted. For example, when the second propulsor 114 is a jet engine thelongitudinal thrust may be adjusted by adjusting a fuel flow rate, anozzle size, a bypass ratio, thrust bleeding, or thrust vectoring.

In another implementation for tilt-rotor aircraft (another particulartype of aircraft 100), the pilot controls the aircraft 100 similar to aconventional helicopter or a conventional VTOL aircraft in the firstregime. Upon a change to the second regime, the first and secondpropulsors 112, 114 tilt forward, and the pilot controls the aircraft100 similar to a conventional airplane. When the first and secondpropulsors 112, 114 are tilt-rotors (known as proprotors), the propulsoractuator 116 receives a command to adjust a nacelle angle of thetilt-rotors (e.g., a pitch angle or orientation of the proprotors). Insome implementations, the command indicates the adjustment to thenacelle angle in an angle value (e.g., radians). In otherimplementations, the command indicates the adjustment to the nacelleangle in an angular rate (e.g., rate of change of the nacelle angle). Bya adjusting a direction of the thrust vector generated by the propulsors112, 114, the longitudinal thrust and speed of the aircraft 100 iscontrolled to operate the aircraft in a desired state.

The control system 104 of the aircraft 100 reduces pilot workload,enables the aircraft 100 to fly more efficiently, and enables theaircraft 100 to have increased capabilities. For example, by replacing aconventional control system of a particular aircraft 100 with thecontrol system 104, the particular aircraft 100 can increase sprinttimes by 30 percent and produce more consistent sprint times.

FIGS. 2A-2D illustrate the effects caused by inceptors of the aircraft100. As illustrated in FIGS. 2A-2D, the aircraft 100 is a compoundhelicopter. FIG. 2A illustrates force (thrust) changes caused byadjusting propulsor collective blade pitch angle. FIG. 2B illustratesforce changes caused by longitudinal (fore and aft) adjustments of themain rotor (e.g., the first propulsor 112) via the cyclic (e.g., theinceptor device 122). FIG. 2C illustrates force changes caused byelevator pitch angle adjustments via the elevator (e.g., a particularcontrol surface 128). FIG. 2D illustrates force changes caused by mainrotor collective blade pitch angle adjustments via the collective.

FIGS. 3A-3D illustrate the effects caused by inceptors of the aircraft100. As illustrated in FIGS. 3A-3D, the aircraft 100 is a tilt-rotoraircraft. FIG. 3A illustrates force changes caused by a nacelle angle ornacelle angular rate of nacelles of the proprotors (e.g., the propulsors112, 114). FIG. 3B illustrates force changes caused by longitudinal(fore and aft) adjustments of the proprotors via the cyclic. FIG. 3Cillustrates force changes caused by elevator pitch angle adjustments viathe cyclic. FIG. 3D illustrates force changes caused by proprotorcollective blade pitch angle adjustments via the collective. AlthoughFIG. 3D illustrated the force changes in the vertical direction,adjusting the collective blade pitch adjusts the thrust of theproprotors along a direction of the proprotors, which may be adjusted asshown in FIGS. 3A and 3B. In FIG. 3A the proprotors are physicallyadjusted in the pitch axis by adjusting a nacelle angle or nacelleangular rate, as compared to FIG. 2A where the collective blade pitchangle of the second propulsor 114, such as a propeller, is adjusted.

FIG. 4 is a circuit diagram 400 that illustrates an example of thecontrol circuitry 130 of FIG. 1. The circuit diagram 400 can be used tocontrol the aircraft 100 of FIG. 1, such as compound helicopters ortilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D. The controlcircuitry 130 of FIG. 4 includes pitch trim prediction circuitry 402,propulsor trim prediction circuitry 404, processing circuitry 406 and408, a control surface actuator 410, the control surfaces 128, thepropulsor actuator 116, and the propulsors 112, 114.

The pitch trim prediction circuitry 402 is configured to generate apredicted pitch attitude trim value 442. The predicted pitch attitudetrim value 442 is generated at least based on an aircraft velocity 422and a pitch attitude 424. The aircraft velocity 422 includes a verticalvelocity, a horizontal velocity, or a combined velocity. For example,the aircraft velocity 422 may include or correspond to a vertical speedindicated (VSI), an airspeed, a ground speed, a true speed, or anothervelocity. The pitch attitude 424 (e.g., an input pitch attitude of theaircraft) includes or corresponds to a measured pitch attitude generatedby data from one or more sensors, such as the sensors 132 of FIG. 1 or afeedback pitch attitude. Examples of generation of the predicted pitchattitude trim value 442 are described further with respect to FIGS. 5-8.

The predicted pitch attitude trim value 442 is or indicates an estimatedpitch attitude of the aircraft 100 for an altitude hold state of theaircraft 100 or a vertical velocity hold state of the aircraft 100. Thealtitude hold state is a state where the aircraft 100 operates in levelflight, such as when the vertical velocity (e.g., the vertical speedindicated (VSI)) is zero. The vertical velocity hold state is a statewhere the aircraft 100 operates with a fixed climb rate or a fixeddescent rate (e.g., a fixed acceleration) or operates with a fixed,non-zero vertical velocity, such as when the vertical velocity isconstant (no vertical velocity acceleration) and the altitude of theaircraft 100 is changing.

The propulsor trim prediction circuitry 404 is configured to generate apredicted propulsor trim value 444. The predicted propulsor trim value444 is generated at least based on a first input (e.g., a thrust input412 as illustrated in FIG. 4), the aircraft velocity 422, and apropulsor feedback value 432. The predicted propulsor trim value 444 mayinclude or correspond to a predicted propulsor collective blade pitchtrim value 562 of FIG. 5 or a predicted proprotor nacelle angle trimvalue 662 of FIG. 6. The propulsor feedback value 432 may include orcorrespond to a previous value of the propulsor command 452 or ameasured propulsor value corresponding to the propulsor command 452,such as a measured collective blade pitch angle or a measured nacelleangle.

The predicted propulsor collective blade pitch trim value 562 is anestimated value of a propulsor collective blade pitch angle setting thatproduces a magnitude of propulsor thrust (e.g., longitudinal thrust) fora speed hold state of the aircraft 100 by adjusting a magnitude of thethrust generated by the propulsor. In some implementations, thepredicted propulsor collective blade pitch trim value 562 is a negativevalue (indicating thrust in the opposite or forward direction) andcauses a reduction in collective blade pitch angle of the propulsor.Reducing the collective blade pitch angle of the propulsor causes thepropulsor to reduce thrust or to generate thrust in the oppositedirection.

The predicted proprotor nacelle angle trim value 662 is an estimatedvalue of a nacelle angle setting of a nacelle of a proprotor thatproduces a magnitude of propulsor thrust (e.g., longitudinal thrust) fora speed hold state of the aircraft 100 by adjusting a direction of thethrust generated by the proprotor. In other implementations, thepredicted propulsor trim value 444 includes or corresponds to othervalues of thrust effectors, as described with reference to FIG. 1.

The processing circuitry 406 is configured to generate the propulsorcommand 452 based on the predicted pitch attitude trim value 442 and thepredicted propulsor trim value 444. The processing circuitry 406 isconfigured to output the propulsor command 452 to one or more propulsoractuators 116. Example configurations of the processing circuitry 406are described with reference to FIGS. 5-15.

The propulsor command 452 is configured to adjust a thrust effector of apropulsor, such as the propulsor actuator 116 of FIG. 1. For example,the propulsor command 452 is configured to adjust a magnitude of thrustgenerated by the propulsor by commanding the propulsor actuator 116 toadjust the magnitude of the thrust directly or indirectly (e.g., byadjusting a direction of the thrust). The propulsor command 452 isconfigured to cause the aircraft 100 to operate in an airspeed holdstate (e.g., a constant speed or no acceleration state) or anacceleration hold state (e.g., constant rate of change ofspeed/velocity). The propulsor command 452 may include or correspond toa propulsor collective blade pitch angle command, a proprotor nacelleangle command, or a command that adjusts a magnitude or a direction ofthe thrust generated by a propulsor.

The processing circuitry 408 is configured to generate a pitch attitudecommand 454 (e.g., an aircraft pitch attitude command) based on thepredicted pitch attitude trim value 442 and one or more other inputs,such as a second pilot input 414 (e.g., a pitch maneuver input asillustrated in FIG. 4). The processing circuitry 408 is configured tooutput the pitch attitude command 454 to one or more control surfaceactuators 410. Example configurations of the processing circuitry 408are described with reference to FIGS. 5-15. The pitch attitude command454 is configured to adjust a position of one or more control surfaces128 of the aircraft 100. For example, the pitch attitude command 454 isconfigured to adjust the position of a particular control surface 128 bycommanding a particular control surface actuator 410 to adjust theposition of the particular control surface 128. The pitch attitudecommand 454 is configured to cause the aircraft to operate in thealtitude hold state or the vertical velocity hold state. To illustrate,adjusting the position of the particular control surface 128 trims theaircraft to achieve and maintain the altitude hold state or the verticalvelocity hold state.

The one or more control surface actuators 410 are configured to adjustone or more of the control surfaces 128 of the aircraft 100 based on thepitch attitude command 454. The one or more control surface actuators410 may include or correspond to one or more elevator actuators,horizontal stabilizer actuators, spoiler actuators, aileron actuators,rudder actuators, flap actuators, slat actuators, thrust vectoringactuators, etc. The control surface actuators 410 may include pitchactuators (e.g., pitch actuators 520 of FIG. 5), yaw actuators, rollactuators, or a combination thereof.

The propulsor actuator 116 is configured to adjust one or more of thepropulsors 112, 114 of the aircraft 100 based on the propulsor command452. In some implementations, multiple propulsor actuators 116 areconfigured to adjust one or more of the propulsors 112, 114 of theaircraft 100 based on the propulsor command 452. The propulsor actuator116 may include or correspond to a collective blade pitch actuator, anacelle angle actuator, a nacelle angular rate actuator, a nozzle sizeactuator, a nozzle direction actuator, a fuel flow rate actuator, etc.,or a combination thereof.

One or more of the propulsors 112, 114, the propulsor actuator 116, thecontrol surfaces 128, or the control surface actuators 410 correspond toflight control effectors 460 of the aircraft 100. The flight controleffectors 460 are configured to adjust properties (e.g., thrust, speed,acceleration, and attitude, such as pitch, roll and yaw) of the aircraft100 to control states of the aircraft 100.

In other implementations, the control circuitry 130 includes one of thepitch trim prediction circuitry 402 or the propulsor trim predictioncircuitry 404. In implementations where the control circuitry 130includes the pitch trim prediction circuitry 402, the control circuitry130 can generate the commands 452, 454 based on the predicted pitchattitude trim value 442 generated by the pitch trim prediction circuitry402 and independent of a predicted propulsor trim value 444. Inimplementations where the control circuitry 130 includes the propulsortrim prediction circuitry 404, the control circuitry 130 can generatethe propulsor command 452 based on the predicted propulsor trim value444 generated by the propulsor trim prediction circuitry 404 andindependent of a predicted pitch attitude trim value 442. Additionally,the control circuitry 130 can generate the propulsor command 452 (i.e.,the predicted propulsor trim value 444) and the pitch attitude command454 independent of pilot or autopilot inputs, i.e., independent of theinputs 412 and 414. For example, the control circuitry 130 can maintaina particular target state independent of pilot input in response to achange in the aircraft velocity 422 or the pitch attitude 424 (e.g.,from a gust of wind, a change in ambient temperature, a change inambient pressure, etc.). Operation of various embodiments of FIG. 4 arediscussed with reference in FIGS. 5 and 6.

FIG. 5 is a circuit diagram 500 that illustrates an example of thecontrol circuitry 130 of FIG. 1. The circuit diagram 500 can be used tocontrol the aircraft 100 of FIG. 1, such as a compound helicopterillustrated in FIG. 2A.

The control circuitry 130 of FIG. 5 includes the pitch trim predictioncircuitry 402, the propulsor trim prediction circuitry 404, a regimerecognition circuitry 502, an integrator 504, a pitch command model 506,a speed select circuitry 508 (e.g., high speed mode circuitry), anacceleration command circuitry 510 (e.g., low speed mode circuitry), anacceleration controller 512, a pitch controller 514, a combiner 516, apropulsor limiting circuitry 518, a first switch 528, and a secondswitch 530. In other implementations, the control circuitry 130 includesadditional components or fewer components than illustrated in FIG. 5.

The pitch trim prediction circuitry 402 and the propulsor trimprediction circuitry 404 were described with reference to FIG. 4 and aredescribed further with reference to FIGS. 7, 8, and 13. As illustratedin FIG. 5, the pitch trim prediction circuitry 402 is configured toreceive a vertical velocity 534 (e.g., a vertical speed indicated (VSI))and to generate the predicted pitch attitude trim value 442 based on thevertical velocity 534. The pitch trim prediction circuitry 402 isconfigured to output the predicted pitch attitude trim value 442 to thefirst switch 528.

In the example illustrated in FIG. 5, the propulsor trim predictioncircuitry 404 generates a particular type of predicted propulsor trimvalue 444, i.e., the predicted propulsor collective blade pitch trimvalue 562. The propulsor trim prediction circuitry 404 is configured togenerate the predicted propulsor collective blade pitch trim value 562based on the aircraft velocity 422, a selected pitch attitude trim value546, an aircraft pitch attitude command 550, and a propulsor collectiveblade pitch feedback 566. The propulsor collective blade pitch feedback566 indicates a propulsor collective blade pitch angle value and mayinclude or correspond to a previous value of the propulsor command 452(e.g., a propulsor collective blade pitch angle command) or a measuredpropulsor collective blade pitch angle.

During operation, the regime recognition circuitry 502 receives an inputand is configured to output a signal that controls one or more switches,such as the first and second switches 528, 530. As illustrated in FIG.5, the regime recognition circuitry 502 receives the aircraft velocity422 and outputs a regime signal 540 to the first and second switches528, 530. For example, the regime recognition circuitry 502 generatesthe regime signal 540 by comparing the aircraft velocity 422 to a speedthreshold.

The regime signal 540 indicates a current regime of the aircraft 100 andis configured to control operation of the first and second switches 528,530. As illustrated in FIG. 5, the regime signal 540 indicates one oftwo regimes and controls a single-pole, dual-throw switch, in otherimplementations, the regime signal 540 indicates more than two regimesor multiple regime signals 540 are used to control the first and secondswitches 528, 530. In such implementations, the switches 528, 530include additional poles or throws.

The thrust inceptor 522 receives a pilot input (e.g., a thumbwheel inputor a collective input) or an autopilot input, and generates a v-dotcommand signal 542, where v-dot indicates acceleration or rate of changeof speed/velocity. For example, the thrust inceptor 522 (or othercircuitry) performs a table lookup using the thumbwheel input togenerate the v-dot command signal 542. The thrust inceptor 522 outputsthe v-dot command signal 542 to the speed select circuitry 508 and tothe acceleration command circuitry 510.

The pitch trim inceptor 524 receives another pilot input (e.g., a beepswitch input) or another autopilot input and generates a pitch triminput command signal. For example, the pitch trim inceptor 524 (or othercircuitry) performs a table lookup based on the beep switch input togenerate the pitch trim input command signal. The beep switch inputindicates a deviation from a zero value or a new reference value. As anillustrative, non-limiting example, the beep switch input indicatesnegative 1 degree nose down. The pitch trim inceptor 524 outputs thepitch trim input command signal to the integrator 504. The integrator504 receives the pitch trim input command signal from the pitch triminceptor 524 and integrates the pitch trim input command signal togenerate a commanded pitch attitude trim value 544. The integrator 504outputs the commanded pitch attitude trim value 544 to the first switch528.

The first switch 528 is configured to output the predicted pitchattitude trim value 442 or the commanded pitch attitude trim value 544based on the regime signal 540. Responsive to the regime signal 540indicating a first regime (e.g., a low speed mode), the first switch 528outputs the commanded pitch attitude trim value 544 as the selectedpitch attitude trim value 546. Responsive to the regime signal 540indicating a second regime (e.g., a high speed mode), the first switch528 outputs the predicted pitch attitude trim value 442 as the selectedpitch attitude trim value 546. The first switch 528 outputs the selectedpitch attitude trim value 546 to the pitch command model 506, to theacceleration command circuitry 510, and to the propulsor trim predictioncircuitry 404. Thus, in FIG. 5, the control circuitry 130 only uses thepredicted pitch attitude trim value 442 in a particular regime (e.g., ahigh speed mode).

The pitch control inceptor 526 receives another pilot input (e.g., acyclic fore and aft input) or another autopilot input and generates apitch attitude input signal 548. For example, the pitch control inceptor526 (or other circuitry) performs a table lookup using the cyclic inputto generate the pitch attitude input signal 548. The pitch controlinceptor 526 outputs the pitch attitude input signal 548 to theintegrator 504.

The pitch command model 506 generates the aircraft pitch attitudecommand 550 based on the selected pitch attitude trim value 546 and thepitch attitude input signal 548. The pitch command model 506 outputs theaircraft pitch attitude command 550 to the acceleration commandcircuitry 510, to the propulsor trim prediction circuitry 404, and tothe pitch controller 514.

The pitch controller 514 (e.g., a pitch attitude controller) generatesone or more control surface pitch commands 558 based on the aircraftpitch attitude command 550 and outputs the one or more control surfacepitch commands 558 to one or more of the pitch actuators 520. The one ormore pitch actuators 520 (e.g., pitch moment control actuators) adjust aposition of one or more control surfaces 128 of the aircraft 100 basedon the one or more control surface pitch commands 558.

The acceleration command circuitry 510 generates an acceleration commandmode acceleration command 554 based on the selected pitch attitude trimvalue 546 and the aircraft pitch attitude command 550 and outputs theacceleration command mode acceleration command 554 to the second switch530. The speed select circuitry 508 generates a speed select modeacceleration command 552 based on the v-dot command signal 542 and thepredicted pitch attitude trim value 442 and outputs the speed selectmode acceleration command 552 to the second switch 530. The secondswitch 530 outputs the speed select mode acceleration command 552 or theacceleration command mode acceleration command 554 to the accelerationcontroller 512 as a selected acceleration command 556 based on theregime signal 540.

The acceleration controller 512 generates a delta propulsor collectiveblade pitch command 560 based on the selected acceleration command 556and the aircraft velocity 422 and outputs the delta propulsor collectiveblade pitch command 560 to the combiner 516. The delta propulsorcollective blade pitch command 560 includes or corresponds to apropulsor maneuver command. To illustrate, the delta propulsorcollective blade pitch command 560 indicates a value of collective bladepitch to achieve an amount of thrust indicated by a thrust input to thethrust inceptor 522 (and indicated by the v-dot command signal 542 whichis generated based on the thrust input, such as the thrust input 412 ofFIG. 4).

The propulsor trim prediction circuitry 404 generates the predictedpropulsor collective blade pitch trim value 562 based on the aircraftpitch attitude command 550, the aircraft velocity 422, and the propulsorcollective blade pitch feedback 566. The predicted propulsor collectiveblade pitch trim value 562 corresponds to a particular type of thepredicted propulsor trim value 444. The propulsor trim predictioncircuitry 404 outputs the predicted propulsor collective blade pitchtrim value 562 to the combiner 516.

The combiner 516 generates a combined propulsor command 564 (combinedpropulsor collective blade pitch command) based on combining (adding)the delta propulsor collective blade pitch command 560 and the predictedpropulsor collective blade pitch trim value 562. The combiner 516outputs the combined propulsor command 564 to the propulsor limitingcircuitry 518. The propulsor limiting circuitry 518 generates thepropulsor command 452 (a propulsor collective blade pitch angle command)based on the combined propulsor command 564 and one or more limits(e.g., thresholds). As illustrated in FIG. 5, the propulsor limitingcircuitry 518 limits the combined propulsor command 564 based onactuator limits 572, engine limits 574 (e.g., engine horsepower (hp)limits), and propulsor limits 576 (e.g., propulsor hp limits) togenerate the propulsor command 452. The propulsor limiting circuitry 518outputs the propulsor command 452 to the propulsor actuator 116 andoutputs the propulsor command 452 to the propulsor trim predictioncircuitry 404 as the propulsor collective blade pitch feedback 566. Thepropulsor actuator 116 adjusts one or more of the propulsors 112, 114based on the propulsor command 452.

FIG. 6 is a circuit diagram 600 that illustrates another example of thecircuit of FIG. 5 for a propulsor with nacelle angle control. Thecircuit diagram 600 can be used to control the aircraft 100 of FIG. 1,such as a tilt-rotor aircraft illustrated in FIGS. 3A-3D. As compared tothe circuit diagram 500 of FIG. 5, which controls a collective bladepitch angle of a propulsor, the circuit diagram 600 of FIG. 6 isconfigured to control a nacelle angle of proprotors of a tilt-rotoraircraft.

The control circuitry 130 of FIG. 6 includes similar components andoperates similar to the control circuitry 130 of FIG. 5. However, inFIG. 6, the control circuitry 130 is configured to generate a predictedproprotor nacelle angle trim value 662 and a proprotor nacelle anglecommand 652 (as compared to the predicted propulsor collective bladepitch trim value 562 and corresponding command). The nacelle anglecommand 652 is a particular type of the propulsor command 452 of FIG. 4.The control circuitry 130 of FIG. 6 has been simplified and somecomponents and signal of the control circuitry 130 of FIG. 5 are notillustrated for clarity.

Additionally, FIG. 6 illustrates the control circuitry 130 including agravity compensator. The gravity compensator includes a first sinefunction circuitry 612, a second sine function circuitry 614, a combiner616, and a gravity multiplier 618. The gravity compensator is configuredto generate a gravity compensated longitudinal acceleration signal 622(g*sin(Δθ)). In other implementations, such as in the control circuitry130 of FIG. 5, the acceleration command circuitry 510, the propulsortrim prediction circuitry 404, or both include gravity compensators.

During operation, the first sine function circuitry 612 receives theselected pitch attitude trim value 546 and applies the sine function tothe selected pitch attitude trim value 546 to generate the sine of theselected pitch attitude trim value 546. The second sine functioncircuitry 614 receives the aircraft pitch attitude command 550 andapplies the sine function to the aircraft pitch attitude command 550 togenerate the sine of the aircraft pitch attitude command 550. The firstsine function circuitry 612 and the second sine function circuitry 614provide the sine of the selected pitch attitude trim value 546 and thesine of the aircraft pitch attitude command 550 to the combiner 616. Thecombiner 616 generates a longitudinal acceleration signal by subtractingthe sine of the selected pitch attitude trim value 546 from the sine ofthe aircraft pitch attitude command 550. The combiner 616 outputs thelongitudinal acceleration signal to the gravity multiplier 618. Thegravity multiplier 618 multiplies the longitudinal acceleration signalby gravity (e.g., a gravitational constant or an experienced gravityvalue) to generate a gravity compensated longitudinal accelerationsignal 622. The gravity multiplier 618 outputs the gravity compensatedlongitudinal acceleration signal 622 to acceleration command circuitry510 and to the propulsor trim prediction circuitry 404. The gravitycompensated longitudinal acceleration signal 622 indicates a pitchattitude deviation from a reference pitch attitude (e.g., a deviationfrom the aircraft pitch attitude command 550 by the predicted pitchattitude trim value 442 or the commanded pitch attitude trim value 544of the selected pitch attitude trim value 546).

In the example illustrated in FIG. 6, the propulsor trim predictioncircuitry 404 generates a predicted proprotor nacelle angle trim value662 based on a nacelle angle feedback 664 and the gravity compensatedlongitudinal acceleration signal 622. The nacelle angle feedback 664indicates a nacelle angle value and may include or correspond to aprevious value of the nacelle angle command 652 or a measured nacelleangle (e.g., indicated by sensor data). The propulsor trim predictioncircuitry 404 outputs the predicted proprotor nacelle angle trim value662 to the combiner 516.

In the example illustrated in FIG. 6, the acceleration controller 512generates a delta nacelle angle command 660 (e.g., a delta proprotornacelle value) and outputs the delta nacelle angle command 660 to thecombiner 516. The combiner 516 generates the combined propulsor command564 (a combined proprotor nacelle command) based on the delta nacelleangle command 660 and the predicted proprotor nacelle angle trim value662. The propulsor limiting circuitry 518 generates the nacelle anglecommand 652 by limiting the combined propulsor command 564. AlthoughFIGS. 5 and 6 illustrate control circuitry 130 for two types ofaircraft, the control circuitry 130 can be modified to control any typeof longitudinal thruster for any compound aircraft configuration.

FIG. 7 is a circuit diagram 700 that illustrates an example of a circuitfor predicting a pitch trim value, such as the pitch trim predictioncircuitry 402 of FIGS. 4-6. The circuit diagram 700 can be used tocontrol the aircraft 100 of FIG. 1, such as compound helicopters ortilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D. As illustratedin FIG. 7, the circuit diagram 700 is illustrated as pitch trimprediction circuitry 402 for a compound helicopter. The circuit diagram700 of FIG. 7 (along with circuit diagrams 800-1400 of FIGS. 8-14) isillustrated in a frequency domain. In FIGS. 8-14, “tau” (“τ” or“τ_filt”) represents a time constant (e.g., a trim prediction timeconstant) of a filter (corresponding to an RC filter time constant in atime domain), and “s” represents a Laplace variable. Additionally, anglevalues are indicated in radians (RAD) and velocity values are indicatedin feet per second (FPS).

As explained above, the pitch trim prediction circuitry 402 generatesthe predicted pitch attitude trim value 442 based on the verticalvelocity 534, the aircraft velocity 422, and the pitch attitude 424. Asillustrated in FIG. 7, the pitch trim prediction circuitry 402 isconfigured to generate the predicted pitch attitude trim value 442further based on an inverted vertical velocity 732, an estimatedaircraft pitch attitude 754 (e.g., an a priori estimate of pitchattitude in level flight), and a vertical damping derivative (Zw). Thepropulsor trim prediction circuitry 404 includes an inverter 702, lowpass filters 704, 724, and 726, combiners 706, 716, and 730, multipliers708, and 712, a gain circuitry 710, a vertical damping derivative (Zw)circuitry 714, a constant 718, a comparison circuitry 720, and a divider722. As illustrated in FIG. 7, the low pass filters 704, 724, and 726are first order low pass filters (τs+1). In other implementations, thelow pass filters 704, 724, and 726 may be second order low pass filtersor higher order low pass filters.

During operation, the pitch trim prediction circuitry 402 receives thevertical velocity 534, the aircraft velocity 422, the pitch attitude424, the inverted vertical velocity 732, and the estimated aircraftpitch attitude 754 as inputs. In some implementations, the inverter 702receives the vertical velocity 534 and generates the inverted verticalvelocity 732 based on inverting (e.g., multiplying by negative one,changing a sign bit, etc.) the vertical velocity 534. In otherimplementations, as shown in FIG. 7, the inverter receives a transformedvertical velocity 774 and generates the inverted vertical velocity 732based on inverting (e.g., multiplying by negative one, changing a signbit, etc.) the transformed vertical velocity 774. For example, when thevertical velocity 534 corresponds to VSI, the vertical velocity 534 istransformed from a geodetic coordinate system to a body axis coordinatesystem by a coordinate transformation circuitry 772. To illustrate thevertical velocity 534 is multiplied by a product of secant(θ)*secant(ϕ)to transform the vertical velocity 534 into the transformed verticalvelocity 774, where theta represents a pitch angle of the aircraft andphi represents a yaw angle of the aircraft.

The inverted vertical velocity 732 represents a component of thevertical velocity 534 or the transformed vertical velocity 774, such asa negative component thereof, in a body axis. The inverter 702 providesthe inverted vertical velocity 732 to the first low pass filter 704 andto the first combiner 706. The first low pass filter 704 generates a lowpass filtered vertical velocity signal 734 by low pass filtering theinverted vertical velocity 732. The first low pass filter 704 providesthe low pass filtered vertical velocity signal 734 to the first combiner706 and to the first multiplier 708.

The first combiner 706 generates a high pass filtered vertical velocitysignal 736 based on subtracting the low pass filtered vertical velocitysignal 734 from the inverted vertical velocity 732. The first low passfilter 704 and the first combiner 706 act in conjunction to high passfilter the inverted vertical velocity 732 to generate the high passfiltered vertical velocity signal 736. The first combiner 706 providesthe high pass filtered vertical velocity signal 736 to the gaincircuitry 710. The gain circuitry 710 multiplies the high pass filteredvertical velocity signal 736 by a gain to generate a filtered verticalacceleration 738 and provides the filtered vertical acceleration 738 tothe second multiplier 712. As illustrated in FIG. 7, the gain circuitry710 (e.g., an amplifier) applies a gain of an inverse of the timeconstant tau (1/τ) of a trim prediction lag filter (e.g., the 704 and706). To illustrate, the time constant is in seconds, so multiplying avelocity by an inverse of the time constant generates an acceleration.

The comparison circuitry 720 receives the aircraft velocity 422 and anairspeed threshold 760 and outputs the larger of the aircraft velocity422 and the airspeed threshold 760 as a selected speed 746 to thedivider 722. The divider 722 divides the constant 718 (e.g., 1 asillustrated in FIG. 7) by the selected speed 746 to generate an inverseselected speed 748. The divider 722 outputs the inverse selected speed748 to the first and second multipliers 708, 712. The first multiplier708 multiplies the low pass filtered vertical velocity signal 734 andthe inverse selected speed 748 to generate a vertical flight path anglesignal 744 and provides the vertical flight path angle signal 744 to thesecond combiner 716.

The vertical flight path angle signal 744 represents the filtered pitchattitude change to rotate the aircraft (e.g., a fuselage thereof) suchthat a body-axis component of geodetic vertical speed becomes zero(e.g., when the vertical velocity 534 includes or corresponds to VSI).Multiplying the low pass filtered vertical velocity signal 734 by theinverse selected speed 748 effectively applies a small angleapproximation of a tangent function. Thus, the vertical flight pathangle signal 744 represents a small angle approximation of the verticalflight path angle that converges to an exact value of the verticalflight path angle as the error (or difference) between the desired andactual vertical flight path angles approaches zero. FIGS. 7 and 8illustrate an altitude hold (e.g., a zero vertical velocity targetvertical state). For a vertical speed hold state the input to thecoordinate transformation circuitry 772, which is the vertical velocity534 for an altitude hold state, is a signal (e.g., a velocity differencesignal) equal to a desired or input value of vertical speed subtractedfrom the vertical velocity 534. While there is no error signal for atarget vertical state in FIG. 7, there is an implied zero vertical speedreference signal, i.e., the inverted vertical velocity 732. Toillustrate, applying a gain of negative one is implicitly a target stateof zero minus the vertical velocity 534 or the transformed verticalvelocity 774. Thus, the inverted vertical velocity 732 represents a zerovertical speed reference signal.

The second multiplier 712 multiplies the filtered vertical acceleration738 and the inverse selected speed 748 to generate a time rate of changeof flight path angle signal 740 and provides the time rate of change offlight path angle signal 740 to the vertical damping derivative (Zw)circuitry 714. The vertical damping derivative (Zw) circuitry 714divides the time rate of change of flight path angle signal 740 by thevertical damping derivative (Zw) to generate a first signal 742 (e.g., atime rate of change of flight path angle signal divided by the verticaldamping derivative (Zw)) and provides the first signal 742 to the secondcombiner 716.

The second combiner 716 generates a filtered aircraft trim pitchdeviation signal 750 based on adding (combining) the first signal 742and the vertical flight path angle signal 744. The second combiner 716provides the filtered aircraft trim pitch deviation signal 750 to thefourth combiner 730. In implementations in which the pitch trimprediction circuitry 402 receives the vertical velocity 534, theaircraft velocity 422 indicates a horizontal velocity, such as groundspeed, air speed, true speed. Thus, the pitch trim prediction circuitry402 generates the filtered aircraft trim pitch deviation signal 750based on vertical and horizontal speeds of the aircraft, i.e., thevelocities 422, 534.

The second low pass filter 724 receives the pitch attitude 424 andgenerates a low pass filtered pitch attitude signal 752 by low passfiltering the pitch attitude 424. The second low pass filter 724provides the low pass filtered pitch attitude signal 752 to the fourthcombiner 730.

The third low pass filter 726 receives the estimated aircraft pitchattitude 754 and generates a low pass filtered estimated pitch attitude756 by low pass filtering the estimated aircraft pitch attitude 754. Thethird low pass filter 726 provides the low pass filtered estimated pitchattitude 756 to the third combiner 728. The third combiner 728 generatesa high pass filtered estimated pitch attitude 758 based on subtractingthe low pass filtered estimated pitch attitude 756 from the estimatedaircraft pitch attitude 754. The third low pass filter 726 and the thirdcombiner 728 act in conjunction to high pass filter the estimatedaircraft pitch attitude 754 to generate the high pass filtered estimatedpitch attitude 758. The third combiner 728 provides the high passfiltered estimated pitch attitude 758 to the fourth combiner 730. Thehigh pass filtered estimated pitch attitude 758 represents a feedforwarda priori estimate of the trim pitch attitude.

The fourth combiner 730 (e.g., output circuitry) generates the predictedpitch attitude trim value 442 based on adding (combining) the filteredaircraft trim pitch deviation signal 750, a low pass filtered pitchattitude signal 752, and the high pass filtered estimated pitch attitude758. The fourth combiner 730 (e.g., the output circuitry) provides thepredicted pitch attitude trim value 442 to the first switch 528, asdescribed with reference to FIGS. 5 and 6. In other implementations, thepitch trim prediction circuitry 402 generates the predicted pitchattitude trim value 442 independent of the estimated aircraft pitchattitude 754. The signals 742, 744, 752, and 758 each have acomplementary dynamic response such that the signals 742, 744, 752, and758 may be summed to estimate the trim pitch attitude for the targetvertical state.

FIG. 8 is a circuit diagram 800 that illustrates another example of acircuit for predicting a pitch trim value, such as the pitch trimprediction circuitry 402 of FIGS. 4-6. The circuit diagram 800 can beused to control the aircraft 100 of FIG. 1, such as compound helicoptersor tilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D. Asillustrated in FIG. 8, the circuit diagram 800 is illustrated as pitchtrim prediction circuitry 402 for a compound helicopter.

The circuit diagram 800 of FIG. 8 uses a normalized dimensional aircraftvertical force per angle of attack derivative (Zα) instead of using thevertical damping derivative (Zw), as in the circuit diagram 700 of FIG.7. As compared to the vertical damping derivative (Zw), the normalizeddimensional aircraft vertical force per angle of attack derivative (Zα)is more complex and more accurate for a high speed mode, a high speedaircraft, and/or aircraft that operate more like airplanes thanhelicopters. For example, the normalized dimensional aircraft verticalforce per angle of attack derivative (Zα) varies significantly withspeed and multiple the normalized dimensional aircraft vertical forceper angle of attack derivative (Zα) values are used. A particularnormalized dimensional aircraft vertical force per angle of attackderivative (Zα) value may be calculated or retrieved from a table duringoperation.

The vertical damping derivative (Zw) is less complex and is accurateduring normal operating envelopes for helicopters and low speedaircraft. For example, the vertical damping derivative (Zw) of ahelicopter or a low speed aircraft is substantially constant foroperational pitch attitudes and speed. Thus, the vertical dampingderivative (Zw) (i.e., a single value) can be approximated and used fora range of aircraft speeds and such that any errors do not affectoperation of the control circuitry 130.

As explained above, the pitch trim prediction circuitry 402 generatesthe predicted pitch attitude trim value 442 based on the verticalvelocity 534 (e.g., a transformed vertical velocity 774 as shown in FIG.8), the aircraft velocity 422, the pitch attitude 424, and the estimatedaircraft pitch attitude 754. The propulsor trim prediction circuitry 404includes the inverter 702, the low pass filters 704, 724, 726, thecombiners 706, 716, 730, the gain circuitry 710, the comparisoncircuitry 720, a divider 802, and a normalized dimensional aircraftvertical force per angle of attack derivative (Zα) circuitry 804. Thecoordinated transformation circuitry 772 is not shown in FIG. 8.

During operation, the pitch trim prediction circuitry 402 of FIG. 8operates similar to the pitch trim prediction circuitry 402 of FIG. 7.In FIG. 8, the low pass filter 704 provides the low pass filteredvertical velocity signal 734 to a divider 802 (as opposed to the firstmultiplier 708 in FIG. 7). The divider 802 generates the vertical flightpath angle signal 744 based on the selected speed 746 (rather than theinverse selected speed 748 as in the circuit diagram 700 of FIG. 7). Thedivider 802 provides the vertical flight path angle signal 744 to thesecond combiner 716 similar to the first multiplier 708 of FIG. 7.

In contrast to FIG. 7, the gain circuitry 710 outputs the filteredvertical acceleration 738 to the normalized dimensional aircraftvertical force per angle of attack derivative (Zα) circuitry 804, asopposed to the gain circuitry 710 outputting to the filtered verticalacceleration 738 to the second multiplier 712. The normalizeddimensional aircraft vertical force per angle of attack derivative (Zα)circuitry 804 divides the filtered vertical acceleration 738 by thenormalized dimensional aircraft vertical force per angle of attackderivative (Zα) to generate a second signal 812 and provides the signal812 to the second combiner 716. The second combiner 716 generates thefiltered aircraft trim pitch deviation signal 750 based on adding(combining) the normalized signal 812 and the vertical flight path anglesignal 744. The second combiner 716 provides the filtered aircraft trimpitch deviation signal 750 to the fourth combiner 730 similar to FIG. 7.

The fourth combiner 730 generates the predicted pitch attitude trimvalue 442 based on adding (combining) the filtered aircraft trim pitchdeviation signal 750, the low pass filtered pitch attitude signal 752,and the high pass filtered estimated pitch attitude 758. The fourthcombiner 730 provides the predicted pitch attitude trim value 442 to thefirst switch 528, as described with reference to FIGS. 5 and 6. In otherimplementations, the pitch trim prediction circuitry 402 generates thepredicted pitch attitude trim value 442 independent of the estimatedaircraft pitch attitude 754.

The signals 742 and 812 represent the filtered pitch attitude changepredicted to arrest or inhibit vertical acceleration, i.e., to achieve atarget vertical state of zero vertical velocity. The signals aregenerated based on a high-pass filter of the vertical velocity 534. Thepitch trim prediction circuitry 402 uses a priori knowledge (i.e., theestimated aircraft pitch attitude 754) of how pitch attitude affectsvertical acceleration for the particular aircraft. The a prioriknowledge may be expressed as a change in vertical force per change inangle of attack derivative (Zα) (804) or as a change in vertical forceper change in vertical velocity derivative (Zw). In otherimplementations, the priori knowledge may be expressed as another valuein a stability axis, a wind axis, or a body axis.

FIG. 9 is a circuit diagram 900 that illustrates an example of a pitchcommand model, such as the pitch command model 506 of FIGS. 5 and 6. Thecircuit diagram 900 can be used to control the aircraft 100 of FIG. 1,such as compound helicopters or tilt-rotor aircraft illustrated in FIGS.2A-2D and 3A-3D. As illustrated in FIG. 9, the circuit diagram 900 isillustrated as a pitch command model 506 for a compound helicopter.

As explained above, the pitch command model 506 generates the aircraftpitch attitude command 550 based on the selected pitch attitude trimvalue 546 and the pitch attitude input signal 548. The pitch commandmodel 506 includes multiple amplifiers 912, 916, and 922, combiners 914and 920, limiters 918 and 924, and integrators 926 and 930 and includesa coordinate transformation circuitry 928. The pitch command model 506includes two feedback loops, such as an inner feedback loop and an outerfeedback loop.

During operation, the first amplifier 912 (e.g., a stick sensitivityamplifier) receives the pitch attitude input signal 548 and amplifiesthe pitch attitude input signal 548. The first amplifier 912 outputs theamplified pitch attitude input signal 548 to the second combiner 920.The first combiner 914 receives the selected pitch attitude trim value546 and the outer loop feedback signal 948 and generates an aircraftpitch attitude command error signal 932 by subtracting the outer loopfeedback signal 948 from the selected pitch attitude trim value 546. Thefirst combiner 914 outputs the aircraft pitch attitude command errorsignal 932 to the second amplifier 916 (e.g., a command filter attitudeamplifier). The second amplifier 916 generates an aircraft pitch ratecommand signal 934 by applying gain (e.g., a command filter attitudegain k2) to the aircraft pitch attitude command error signal 932. Thesecond amplifier provides the aircraft pitch rate command signal 934 tothe first authority limiter 918.

The first authority limiter 918 generates a limited aircraft pitch ratecommand 936 by limiting the aircraft pitch rate command signal 934. Forexample, the first authority limiter 918 reduces or increases a value ofthe aircraft pitch rate command signal 934 based on one or morethresholds (e.g., authority thresholds). For example, the firstauthority limiter 918 compares the value of the aircraft pitch ratecommand signal 934 to a maximum pitch attitude value (e.g., a firstthreshold) and reduces the value of the aircraft pitch rate commandsignal 934 to the maximum pitch attitude value. In some implementations,the one or more thresholds of the first authority limiter 918 are basedon conditions of the aircraft. To illustrate, the maximum pitch value(e.g., the first threshold) has multiple values (e.g., multiplethresholds or a variable threshold) depending on a speed of the aircraftand a temperature of the ambient air. The first authority limiter 918outputs the limited aircraft pitch rate command 936 to the secondcombiner 920.

The second combiner 920 generates an aircraft pitch rate command errorsignal 940 based on subtracting an inner loop feedback signal 938 from asum of the amplified pitch attitude input signal 548 and the limitedaircraft pitch rate command 936. The second combiner 920 outputs thelimited aircraft pitch rate command error signal 940 to the thirdamplifier 922 (e.g., a command filter rate gain amplifier). The thirdamplifier 922 generates the amplified aircraft pitch rate command errorsignal 940 by applying gain (e.g., a command rate filter gain k1) to theaircraft pitch attitude command error signal 932. The amplified aircraftpitch rate command error signal 940 indicates a pitch angularacceleration demand of the aircraft, such as a quantity of pitch momentto add or subtract based on the pitch attitude input signal 548 from thepitch control inceptor 526. The third amplifier 922 provides theamplified aircraft pitch rate command error signal 940 to the secondauthority limiter 924.

The second authority limiter 924 generates a pitch angular accelerationcommand 942 by limiting the amplified aircraft pitch rate command errorsignal 940. For example, the second authority limiter 924 reduces orincreases a value of the aircraft pitch rate command error signal 940based on one or more thresholds (e.g., authority thresholds). Forexample, the second authority limiter 924 compares the value of theaircraft pitch rate command error signal 940 to a maximum pitch value(e.g., a first threshold) and reduces the value of the aircraft pitchrate command error signal 940 to the maximum pitch value. In someimplementations, the one or more thresholds of the second authoritylimiter 924 are based on conditions of the aircraft. To illustrate, themaximum pitch value (e.g., the first threshold) has multiple values(e.g., multiple thresholds or a variable threshold) depending on a speedof the aircraft 100 and a temperature of the ambient air. The secondauthority limiter 924 outputs the limited aircraft pitch rate command936 to the first integrator 926.

The first integrator 926 integrates the pitch angular accelerationcommand 942 to generate a pitch rate command 944. The first integrator926 provides the pitch rate command 944 to the coordinate transformationcircuitry 928. Additionally, the first integrator 926 provides the pitchrate command 944 as the inner loop feedback signal 938 to the secondcombiner 920. The coordinate transformation circuitry 928 generates apitch attitude rate command 946 by transforming the pitch rate command944 from a first coordinate system to a second coordinate system. Forexample, the coordinate transformation circuitry 928 applies a Body toEuler coordinate transformation to transform the pitch rate command 944into Euler coordinates. The coordinate transformation circuitry 928outputs the pitch attitude rate command 946 to the second integrator930. The second integrator 930 integrates the pitch attitude ratecommand 946 to generate the aircraft pitch attitude command 550.Additionally, the second integrator 930 provides the aircraft pitchattitude command 550 as an outer loop feedback signal 948 to the firstcombiner 914.

FIG. 10 is a circuit diagram 1000 that illustrates an example of acircuit for a high speed mode or a speed select mode, such as the speedselect circuitry 508 of FIGS. 5 and 6. The circuit diagram 1000 can beused to control the aircraft 100 of FIG. 1, such as compound helicoptersor tilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D. In FIG. 10,the circuit diagram 1000 is illustrated as speed select circuitry 508for a compound helicopter.

As explained above, the speed select circuitry 508 generates the speedselect mode acceleration command 552 based on the v-dot command signal542 and the predicted pitch attitude trim value 442. The speed selectcircuitry 508 includes multiple amplifiers 1012, 1026, and 1030 andcombiners 1016 and 1024 and includes a comparison circuitry 1014, aswitch 1018, and an integrator 1022. The speed select circuitry 508includes a feedback loop and one terminal of the switch 1018 is coupledto ground 1020.

The switch 1018 is configured to output a pitch independent accelerationcommand error signal 1034 or a ground signal based on a control signal1036 generated by the comparison circuitry 1014. As illustrated in FIG.10, the switch 1018 is a single-pole, dual-throw switch and for theillustrated comparison condition (i.e., absolute value greater thanzero) of the comparison circuitry 1014, the switch 1018 outputs theground signal when the pitch independent acceleration command errorsignal 1034 is zero.

During operation, the first amplifier 1012 (e.g., an inceptorsensitivity amplifier) receives the v-dot command signal 542 and appliesa first gain (e.g., a function dependent inceptor sensitivity gainF(delta)) to the v-dot command signal 542 to generate a pitchindependent acceleration command 1032. For example, the first amplifier1012 (or other circuitry) performs a table lookup using the v-dotcommand signal 542 to determine the first gain or to determine the pitchindependent acceleration command 1032. As another example, the firstamplifier 1012 (or other circuitry) performs a table lookup usinganother variable (e.g., speed) to determine the first gain, and thefirst amplifier 1012 applies the first gain to the v-dot command signal542 to generate the pitch independent acceleration command 1032.

The first amplifier 1012 outputs the pitch independent accelerationcommand 1032 to the comparison circuitry 1014 and the first combiner1016. The first combiner 1016 generates the pitch independentacceleration command error signal 1034 based on subtracting the washoutsignal 1046 from the pitch independent acceleration command 1032. Thefirst combiner 1016 provides the pitch independent acceleration commanderror signal 1034 to the switch 1018.

The comparison circuitry 1014 generates the control signal 1036 based ona threshold or a comparison condition. As illustrated in FIG. 10, thecomparison circuitry 1014 compares an absolute value of the pitchindependent acceleration command 1032 to zero. When the absolute valueof the pitch independent acceleration command 1032 is greater than zero(e.g., when the pitch independent acceleration command 1032 is not equalto zero), the comparison circuitry 1014 outputs the control signal 1036having a first logical value (e.g., 1). When the absolute value of thepitch independent acceleration command 1032 is not greater than zero(i.e., when the pitch independent acceleration command 1032 is equal tozero), the comparison circuitry 1014 outputs the control signal 1036having a second logical value (e.g., 0). As illustrated in FIG. 10, theabsolute value of the pitch independent acceleration command 1032 isgreater than zero, the comparison circuitry 1014 outputs the controlsignal 1036 having the first logical value (e.g., 1), and the switch1018 outputs the control signal 1036 to the integrator 1022.

As illustrated in FIG. 10, the integrator 1022 integrates the pitchindependent acceleration command error signal 1034 to generate a pitchindependent speed command 1038. The integrator 1022 outputs the pitchindependent speed command 1038 to the second combiner 1024. The secondcombiner 1024 generates a speed error signal 1040 based on subtractingthe aircraft velocity 422 from the pitch independent speed command 1038.The second combiner 1024 provides the speed error signal 1040 to thesecond amplifier 1026 and provides the speed error signal 1040 to thedeadzone limiter 1028 as a speed error feedback signal 1042. The secondamplifier 1026 (e.g., a velocity error amplifier) receives the speederror signal 1040 and applies a second gain (e.g., a velocity errorgain) to the speed error signal 1040 to generate the speed select modeacceleration command 552. The second gain (e.g., the velocity errorgain) may be a constant value (Kverr) or a function dependent value(F(verr)). A function dependent gain enables non-linear control of thespeed select mode acceleration command 552. The second amplifier 1026outputs the speed select mode acceleration command 552 to the secondswitch 530, as illustrated in FIGS. 5 and 6.

The deadzone limiter 1028 generates a limited speed error feedbacksignal 1044 by limiting the speed error feedback signal 1042. Forexample, the deadzone limiter 1028 reduces or increases a value of thelimited speed error feedback signal 1044 based on one or more thresholds(e.g., deadzone thresholds). For example, the deadzone limiter 1028compares the value of the speed error feedback signal 1042 to a maximumvelocity error value (e.g., a first threshold) and reduces the value ofthe speed error feedback signal 1042 to the maximum velocity errorvalue. In some implementations, the one or more thresholds of thedeadzone limiter 1028 are based on conditions of the aircraft 100. Toillustrate, the maximum velocity error value (e.g., the first threshold)has multiple values (e.g., multiple thresholds or a variable threshold)depending on a speed of the aircraft 100 and a temperature of theambient air. The deadzone limiter 1028 outputs the limited speed errorfeedback signal 1044 to the third amplifier 1030.

The third amplifier 1030 (e.g., a washout amplifier) receives thelimited speed error feedback signal 1044 and applies a third gain (e.g.,a washout gain Kwo) to the limited speed error feedback signal 1044 togenerate a washout signal 1046. The third amplifier 1030 outputs thewashout signal 1046 to the first combiner 1016.

FIG. 11 is a circuit diagram 1100 that illustrates an example of acircuit for a low speed mode or an acceleration command mode, such asthe acceleration command circuitry 510 of FIGS. 5 and 6. The circuitdiagram 1100 can be used to control the aircraft 100 of FIG. 1, such ascompound helicopters or tilt-rotor aircraft illustrated in FIGS. 2A-2Dand 3A-3D. In FIG. 11, the circuit diagram 1100 is illustrated asacceleration command circuitry 510 for a compound helicopter.

As explained above, the acceleration command circuitry 510 generates theacceleration command mode acceleration command 554 based on the v-dotcommand signal 542, the selected pitch attitude trim value 546, and theregime signal 540. The acceleration command circuitry 510 includes afirst amplifier 1012, a first combiner 1114, the first sine functioncircuitry 612, the combiner 616, the second sine function circuitry 614,the gravity multiplier 618, and a gravity amplifier 1124.

The gravity amplifier 1124 is configured to amplify the gravitycompensated longitudinal acceleration signal 622 by a gain value togenerate a gravity amplified signal (a gravity compensated command1142). In some implementations, the gain value is a function dependentgain and is generated based on one or more variables, such as velocity,acceleration, pitch attitude, etc. In other implementations, the gain isa constant value or the gain is determined by performing a table lookupto a table including gravity amplifier gain values. For example, thetable indicates the gain is a value of one for small commanded changesin pitch attitude to provide harmony with the lateral axis. The tableindicates a larger gain value (i.e., increases the gain of the gravityamplifier 1124) for larger commanded changes in pitch attitude to avoidlarge pitch attitudes and the commensurate obstruction of pilot sightlines at large pitch attitudes. Additionally, the gain value may be zerosuch that commanded changes in pitch attitude do not command changes invelocity. The table lookup is performed based on a single variable(e.g., speed) or based on multiple variables (e.g., speed and pitchattitude).

During operation, the first amplifier 1012 (e.g., an inceptorsensitivity amplifier) receives the v-dot command signal 542 and appliesa first gain (e.g., a function dependent inceptor sensitivity gainF(delta)) to the v-dot command signal 542 to generate the pitchindependent acceleration command 1032, as described with reference toFIG. 10. The first amplifier 1012 outputs the pitch independentacceleration command 1032 to the first combiner 1114. The first combiner1114 generates the acceleration command mode acceleration command 554based on subtracting the gravity compensated command 1142 from the pitchindependent acceleration command 1032. The first combiner 1114 providesthe acceleration command mode acceleration command 554 to the secondswitch 530 of FIG. 5, as described with reference to FIGS. 5 and 6.

The first sine function circuitry 612 receives the selected pitchattitude trim value 546 and applies the sine function to the selectedpitch attitude trim value 546 to generate a sine of the selected pitchattitude trim value 1134. The second sine function circuitry 614receives the aircraft pitch attitude command 550 and applies the sinefunction to the aircraft pitch attitude command 550 to generate a sineof the aircraft pitch attitude command 1136. The first sine functioncircuitry 612 and the second sine function circuitry 614 provide thesine of the selected pitch attitude trim value 1134 and the sine of theaircraft pitch attitude command 1136 to the combiner 616. The combiner616 generates a longitudinal acceleration signal 1138 by subtracting thesine of the selected pitch attitude trim value 1134 from the sine of theaircraft pitch attitude command 1136. The combiner 616 outputs thelongitudinal acceleration signal 1138 to the gravity multiplier 618. Thegravity multiplier 618 multiplies the longitudinal acceleration signal1138 by gravity (e.g., an acceleration due to gravity constant or anexperienced gravity value) to generate the gravity compensatedlongitudinal acceleration signal 622. The gravity multiplier 618 outputsthe gravity compensated longitudinal acceleration signal 622 to thegravity amplifier 1124.

The gravity amplifier 1124 generates the gravity compensated command1142 based on the gravity compensated longitudinal acceleration signal622. For example, the gravity amplifier 1124 applies a first gain (e.g.,a function dependent gravity amplifier gain F(v-dot)) to the gravitycompensated longitudinal acceleration signal 622 to generate the gravitycompensated command 1142. The first gain (e.g., the function dependentgravity amplifier gain F(v-dot)) may be dependent on one or morefactors, such as velocity, acceleration, attitude, etc. In a particularimplementation, the gravity amplifier 1124 retrieves the first gainvalue from a table. The gravity amplifier 1124 outputs the gravitycompensated command 1142 to the first combiner 1114.

FIG. 12 is a circuit diagram 1200 that illustrates an example of acircuit for generating a delta propulsor command, such as theacceleration controller 512 of FIGS. 5 and 6. The circuit diagram 1200can be used to control the aircraft 100 of FIG. 1, such as compoundhelicopters or tilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D.In FIG. 12, the circuit diagram 1200 is illustrated as an accelerationcontroller 512 for a compound helicopter.

As explained above, the acceleration controller 512 generates the deltapropulsor collective blade pitch command 560 based on the predictedpitch attitude trim value 442 and the selected acceleration command 556.In FIG. 12, the acceleration controller 512 further generates the deltapropulsor collective blade pitch command 560 based on an input v-dot1230 (e.g., measured acceleration or acceleration feedback). Theacceleration controller 512 includes a first combiner 1212, an amplifier1214, an authority limiter 1216, a second combiner 1218, a divider 1220,and a propulsor sensitivity scheduler 1222.

The propulsor sensitivity scheduler 1222 is configured to generate apropulsor sensitivity value 1242 based on the aircraft velocity 422. Forexample, the propulsor sensitivity scheduler 1222 generates thepropulsor sensitivity value 1242 by performing a table lookup using theaircraft velocity 422. As another example, the propulsor sensitivityscheduler 1222 calculates the propulsor sensitivity value 1242 using afunction where the aircraft velocity 422 is an input variable. Otheraircraft or propulsor characteristics can be used as the input variable,such as propeller rotational speed. The propulsor sensitivity value 1242is configured to adjust or convert units of a signal (i.e., performdimensional conversion or dimensional analysis). For example, asillustrated in FIG. 12, the propulsor sensitivity value 1242 convertsthe units of a combined acceleration command 1240 (e.g., having a valueindicating acceleration) into another type of units (e.g., having avalue indicating force in terms of propulsor values, such as force perpropulsor collective blade angle).

During operation, the acceleration controller 512 receives the selectedacceleration command 556 from the second switch 530 of FIG. 5, asdescribed with reference to FIGS. 5 and 6. The selected accelerationcommand 556 is provided to the first combiner 1212 and is provided as afeedforward acceleration command 1232 to second combiner 1218. The firstcombiner 1212 generates an acceleration error signal 1234 by subtractingthe input v-dot 1230 from the selected acceleration command 556 andprovides the acceleration error signal 1234 to the amplifier 1214 (e.g.,a v-dot feedback amplifier). The amplifier 1214 applies a first gain(e.g., a v-dot feedback gain Kvdfb) to the acceleration error signal1234 to generate an acceleration feedback command 1236. The amplifier1214 outputs the acceleration feedback command 1236 to the authoritylimiter 1216 (e.g., authority limiting circuitry).

The authority limiter 1216 generates a limited acceleration feedbackcommand 1238 by limiting the acceleration feedback command 1236. Forexample, the authority limiter 1216 reduces or increases a value of theacceleration feedback command 1236 based on one or more thresholds(e.g., authority thresholds). For example, the authority limiter 1216compares the value of the acceleration feedback command 1236 to amaximum (or minimum) velocity error value (e.g., a first threshold) andreduces the value of the acceleration feedback command 1236 to themaximum velocity error value. In some implementations, the one or morethresholds of the authority limiter 1216 are based on conditions of theaircraft 100. To illustrate, the velocity error value (e.g., the firstthreshold) has multiple values (e.g., multiple thresholds or a variablethreshold) depending on a speed of the aircraft 100 and a temperature ofthe ambient air. The authority limiter 1216 outputs the limitedacceleration feedback command 1238 to the second combiner 1218.

The second combiner 1218 generates the combined acceleration command1240 by adding (combining) the feedforward acceleration command 1232 andthe limited acceleration feedback command 1238 and provides the combinedacceleration command 1240 to the divider 1220. The divider 1220generates the delta propulsor collective blade pitch command 560 bydividing the combined acceleration command 1240 by a propulsorsensitivity value 1242. The propulsor sensitivity scheduler 1222generates the propulsor sensitivity value 1242 based on the aircraftvelocity 422. The divider 1220 outputs the delta propulsor collectiveblade pitch command 560 to the combiner 516 of FIG. 5, as described withreference to FIGS. 5 and 6.

FIG. 13 is a circuit diagram 1300 that illustrates an example of acircuit for predicting a propulsor trim value, such as the propulsortrim prediction circuitry 404 of FIGS. 4-6. The circuit diagram 1300 canbe used to control the aircraft 100 of FIG. 1, such as compoundhelicopters or tilt-rotor aircraft illustrated in FIGS. 2A-2D and 3A-3D.In FIG. 13, the circuit diagram 1300 is illustrated as propulsor trimprediction circuitry 404 for a compound helicopter.

As explained with reference to FIGS. 4-6, the propulsor trim predictioncircuitry 404 is configured to generate the predicted propulsorcollective blade pitch trim value 562 based on the aircraft velocity422, the propulsor collective blade pitch feedback 566, the gravitycompensated longitudinal acceleration signal 622, or a combinationthereof. The propulsor trim prediction circuitry 404 includes low passfilters 1312-1316, combiners 1322-1328, a gain circuitry 1330, a divider1332, and the propulsor sensitivity scheduler 1222. Although, thepropulsor trim prediction circuitry 404 receives the gravity compensatedlongitudinal acceleration signal 622 in FIG. 13, in otherimplementations, the propulsor trim prediction circuitry 404 receivesselected pitch attitude trim value 546 and the aircraft pitch attitudecommand 550, as illustrated in FIG. 5. In such implementations, thepropulsor trim prediction circuitry 404 includes a gravity compensator(e.g., the components 612-618 of FIG. 6) configured to generate thegravity compensated longitudinal acceleration signal 622 based on theselected pitch attitude trim value 546 and the aircraft pitch attitudecommand 550, as described wither reference to FIGS. 6 and 11.

During operation, the propulsor trim prediction circuitry 404 receivesthe aircraft velocity 422, the propulsor collective blade pitch feedback566, and the gravity compensated longitudinal acceleration signal 622.As illustrated in FIG. 13, the first low pass filter 1312 and the firstcombiner 1322 receive the gravity compensated longitudinal accelerationsignal 622. The first low pass filter 1312 filters the gravitycompensated longitudinal acceleration signal 622 to generate a low passfiltered gravity compensated longitudinal acceleration signal 1342. Thefirst combiner 1322 further receives the low pass filtered gravitycompensated longitudinal acceleration signal 1342 and subtracts the lowpass filtered gravity compensated longitudinal acceleration signal 1342from the gravity compensated longitudinal acceleration signal 622 togenerate a high pass filtered gravity compensated longitudinalacceleration signal 1344. The first combiner 1322 outputs the high passfiltered gravity compensated longitudinal acceleration signal 1344 tothe third combiner 1326.

The second low pass filter 1314 and the second combiner 1324 receive theaircraft velocity 422 (or a scalar value, i.e., speed). The second lowpass filter 1314 filters the aircraft velocity 422 to generate a lowpass filtered aircraft velocity 1346. The second combiner 1324 furtherreceives the low pass filtered aircraft velocity 1346 and subtracts thelow pass filtered aircraft velocity 1346 from the aircraft velocity 422to generate a high pass filtered aircraft velocity 1348 (or a high passfiltered speed). The second combiner 1324 outputs the high pass filteredaircraft velocity 1348 to the gain circuitry 1330. The gain circuitry1330 applies a gain to the high pass filtered aircraft velocity 1348 togenerate a filtered acceleration feedback 1350. The filteredacceleration feedback 1350 is a signal that indicates an estimatedacceleration of the aircraft. As illustrated in FIG. 13, the aircraftvelocity 422 is filtered by a first order low pass filter (τs+1) andthen the difference of the aircraft velocity 422 and the low passfiltered aircraft velocity 1346 (i.e., the high pass filtered aircraftvelocity 1348) is multiplied by a gain of the inverse of the trimprediction lag filter time constant (1/τ) to generate the filteredacceleration feedback 1350. Thus, the filtered acceleration feedback1350 has a transfer function of s/(τs+1), where “s” represents aderivative of speed, i.e., acceleration. The gain circuitry 1330 outputsthe filtered acceleration feedback 1350 to the third combiner 1326.

The third combiner 1326 (e.g., intermediary circuitry) subtracts thefiltered acceleration feedback 1350 from the high pass filtered gravitycompensated longitudinal acceleration signal 1344 to generate an offsetvalue 1352 (e.g., a longitudinal acceleration offset value). The thirdcombiner 1326 outputs the offset value 1352 to the divider 1332. Thedivider 1332 divides the offset value 1352 by the propulsor sensitivityvalue 1242 to generate a filtered longitudinal control effector error1356 (e.g., a propulsor force offset value). The filtered longitudinalcontrol effector error 1356 (e.g., a derived propulsor value) indicatesan error based on (e.g., of or from) the pitch attitude deviation interms of an amount of propulsor force or a propulsor setting value.

As explained with reference to FIG. 12, the propulsor sensitivityscheduler 1222 generates the propulsor sensitivity value 1242 (propulsorcontrol derivative or propulsor sensitivity or propulsor controlsensitivity) based on the aircraft velocity 422. For example, asillustrated in FIG. 13, the propulsor sensitivity value 1242 convertsthe units of the offset value 1352 (e.g., having a value indicatingacceleration) into another type of units (e.g., having a valueindicating force in terms of propulsor values, such as force perpropulsor collective blade angle). In some implementations, thepropulsor sensitivity value 1242 includes or corresponds to normalizeddimensional aircraft longitudinal force per propulsor collective bladepitch angle derivative value.

The third low pass filter 1316 receives the propulsor collective bladepitch feedback 566 and low pass filters the propulsor collective bladepitch feedback 566 to generate a low pass filtered propulsor collectiveblade pitch feedback 1358. The third low pass filter 1316 outputs thelow pass filtered propulsor collective blade pitch feedback 1358 to thefourth combiner 1328. The fourth combiner 1328 (e.g., output circuitry)combines the filtered longitudinal control effector error 1356 and thelow pass filtered propulsor collective blade pitch feedback 1358 togenerate the predicted propulsor collective blade pitch trim value 562.The fourth combiner 1328 (e.g., the output circuitry) outputs thepredicted propulsor collective blade pitch trim value 562 to thecombiner 516 of FIG. 5, as described with reference to FIG. 5.

In other implementations, other types of filters or equivalent circuitrymay be used in FIGS. 7-13. For example, a high pass filter may be usedin place of a combiner and a low pass filter. As another example, a highpass filter and combiner may be used in place of a low pass filter togenerate a low pass filtered signal.

FIG. 14 is a circuit diagram 1400 that illustrates an example of acircuit for propulsor limiting, such as the propulsor limiting circuitry518 of FIGS. 5 and 6. The circuit diagram 1400 can be used to controlthe aircraft 100 of FIG. 1, such as compound helicopters or tilt-rotoraircraft illustrated in FIGS. 2A-2D and 3A-3D. In FIG. 14, the circuitdiagram 1400 is illustrated as propulsor limiting circuitry 518 for acompound helicopter.

As explained with reference to FIG. 5, the propulsor limiting circuitry518 is configured to output the propulsor command 452 based on thecombined propulsor command 564. As illustrated in FIG. 14, the propulsorlimiting circuitry 518 includes rate limiting circuitry 1412 andauthority limiting circuitry 1414. The rate limiting circuitry 1412 isconfigured to limit changes in the combined propulsor command 564 basedon characteristics of the propulsor. For example, the rate limitingcircuitry 1412 is configured to limit an amount of change in thecombined propulsor command 564 over time (e.g., a difference between acurrent value and a previous value of the combined propulsor command564). As illustrated in FIG. 14, the rate limiting circuitry 1412generates a rate limited propulsor command 1422 based on the combinedpropulsor command 564. As an illustrative, non-limiting example, therate limiting circuitry 1412 calculates a difference value between aprevious value of the combined propulsor command 564 and a current valueof the combined propulsor command 564. The rate limiting circuitry 1412compares the difference value to one more thresholds (e.g., ratethresholds). Based on the difference value exceeding a first thresholdvalue or failing to a meet a second threshold value, the value of thecombined propulsor command 564 is adjusted to one of the thresholdvalues.

The authority limiting circuitry 1414 is configured to limit changes inthe rate limited propulsor command 1422 (or the combined propulsorcommand 564) based on characteristics of the propulsor and conditions ofthe propulsor and aircraft 100. As illustrated in FIG. 14, the authoritylimiting circuitry 1414 generates the propulsor command 452 based on therate limited propulsor command 1422. For example, the authority limitingcircuitry 1414 is configured to limit a value of the rate limitedpropulsor command 1422. As an illustrative, non-limiting example, theauthority limiting circuitry 1414 compares the value of the combinedpropulsor command 564 or the rate limited propulsor command 1422 to onemore thresholds (e.g., one or more maximum and minimum values. In someimplementations, the thresholds (e.g., authority thresholds) havedifferent values based on current conditions of the aircraft 100.

During operation, the combiner 516 generates the combined propulsorcommand 564 based on the delta propulsor collective blade pitch command560 and the predicted propulsor collective blade pitch trim value 562,as described with reference to FIG. 5. The combiner 516 provides thecombined propulsor command 564 to the rate limiting circuitry 1412 ofthe propulsor limiting circuitry 518. The rate limiting circuitry 1412generates the rate limited propulsor command 1422 based on the combinedpropulsor command 564. The rate limiting circuitry 1412 outputs the ratelimited propulsor command 1422 to the authority limiting circuitry 1414.The authority limiting circuitry 1414 generates the propulsor command452 based on the rate limited propulsor command 1422. The authoritylimiting circuitry 1414 outputs the propulsor command 452 to thepropulsor actuator 116 of FIGS. 1 and 5, as described with reference toFIG. 5.

FIG. 15 illustrates a particular example of a method 1500 forcontrolling an aircraft, such as the aircraft 100 of FIG. 1. In someimplementations, the aircraft is a compound helicopter, a multi-rotoraircraft, a high speed vertical takeoff and landing (VTOL) aircraft, ora combination thereof. The method 1500 may be performed by a controlcircuitry or components thereof, such as the control circuitry 130 ofFIGS. 1 and 4-14.

The method 1500 includes, at 1502, generating a predicted propulsorcollective blade pitch trim value for a target state of the aircraftbased on an aircraft velocity and a pitch attitude deviation from areference. For example, the predicted propulsor collective blade pitchtrim value may include or correspond to the predicted propulsor trimvalue 444 of FIG. 4 or the predicted propulsor collective blade pitchtrim value 562 of FIG. 5. The target state may include, or correspondto, a target horizontal state (e.g., an airspeed hold state or anacceleration hold state). The aircraft velocity may include, orcorrespond to, the aircraft velocity 422 of FIG. 4 or the verticalvelocity 534 of FIG. 5, and the pitch attitude deviation from thereference may include, or correspond to, the aircraft pitch attitudecommand 550 of FIG. 5 or the gravity compensated longitudinalacceleration signal 622 of FIG. 6.

The method 1500 includes, at 1504, adjusting propulsor collective bladepitch angle of a propulsor of the aircraft based on the predictedpropulsor collective blade pitch trim value. For example, the propulsoractuator 116 adjusts a collective blade pitch angle of a propulsor(e.g., a propeller) of the aircraft 100 of FIGS. 2A-2D, as describedwith reference to FIGS. 1, 2A, 4, 5, and 13.

In some implementations, the predicted propulsor collective blade pitchtrim value is an estimated value of a collective blade pitch anglesetting of the propulsor that produces a magnitude of thrust for anairspeed hold state or an acceleration hold state of the aircraft.Additionally or alternatively, the propulsor collective blade pitchangle command is configured to adjust a magnitude of thrust generated bythe propulsor, and the propulsor collective blade pitch angle command isconfigured to cause the aircraft to operate in an airspeed hold state oran acceleration hold state of the aircraft. In a particularimplementation, the propulsor collective blade pitch angle command(e.g., the propulsor command 452) is output by output circuitry thatincludes, or corresponds to, the combiner 516, the propulsor limitingcircuitry 518, or a combination thereof.

In some implementations, the method 1500 further includes generating agravity compensated longitudinal acceleration based on a sine value of aselected aircraft trim pitch attitude value and a sine value of acommanded pitch attitude, where the predicted propulsor collective bladepitch trim value is generated based on the gravity compensatedlongitudinal acceleration. For example, the gravity compensator 612-618generates the gravity compensated longitudinal acceleration signal 622based on the sine value of the selected pitch attitude trim value 1134and the sine of the aircraft pitch attitude command 1136, as describedwith reference to FIGS. 6 and 11. The propulsor trim predictioncircuitry 404 generates the predicted propulsor trim value 444 based onthe gravity compensated longitudinal acceleration signal 622, asdescribed with reference to FIG. 13.

In some implementations, the method 1500 further includes operating in aspeed select mode based on a regime signal. In a particularimplementation, operating in the speed select mode includes generating apitch independent speed command based on a pitch independentacceleration command, where the pitch independent acceleration commandis generated based on one or more pilot inputs. Operating in the speedselect mode also includes generating a speed error signal based onsubtracting the aircraft velocity from the pitch independent speedcommand, and applying a gain function to the speed error signal togenerate a speed select mode acceleration command. Operating in thespeed select mode further includes generating the propulsor collectiveblade pitch angle command based on the speed select mode accelerationcommand and the predicted propulsor collective blade pitch trim value,the propulsor collective blade pitch angle of the propulsor adjustedbased on the propulsor collective blade pitch angle command. Forexample, the speed select circuitry 508 generates the pitch independentacceleration command 1032 by performing a table lookup using the v-dotcommand signal 542 (e.g., thumbwheel inceptor input). The speed selectcircuitry 508 generates the pitch independent acceleration command errorsignal 1034 based on the independent acceleration command 1032 andintegrates the pitch independent acceleration command error signal 1034to generate the pitch independent speed command 1038. The speed selectcircuitry 508 generates the speed error signal 1040 based on subtractingthe aircraft velocity 422 from the pitch independent speed command 1038and applies the velocity error gain F(verr) to the speed error signal1040 to generate the speed select mode acceleration command 552, asdescribed with reference to FIG. 10.

In some implementations, the method 1500 further includes operating inan acceleration command mode based on a regime signal. In a particularimplementation, operating in the acceleration command mode includesgenerating an acceleration command mode acceleration command based on apitch independent acceleration command and a gravity amplifiedlongitudinal acceleration command, where the pitch independentacceleration command is generated based on one or more pilot inputs.Operating in the acceleration command mode further includes generatingthe propulsor collective blade pitch angle command based on theacceleration command mode acceleration command and the predictedpropulsor collective blade pitch trim value, the propulsor collectiveblade pitch angle of the propulsor adjusted based on the propulsorcollective blade pitch angle command. For example, the accelerationcommand circuitry 510 (e.g., the acceleration command mode circuitry)generates the pitch independent acceleration command 1032 by performinga table lookup using the v-dot command signal 542 (e.g., thumbwheelinceptor input). The acceleration command circuitry 510 generates theacceleration command mode acceleration command 554 based on combiningthe pitch independent acceleration command 1032 and the gravitycompensated command 1142 generated by the gravity amplifier 1124.

FIG. 16 illustrates a particular example of a method 1600 forcontrolling an aircraft, such as the aircraft 100 of FIG. 1. The method1600 may be performed by a control circuitry or components thereof, suchas the control circuitry 130 of FIGS. 1 and 4-14.

The method 1600 includes, at 1602, generating a predicted proprotornacelle trim value for a target state of the aircraft based on anaircraft velocity and a pitch attitude deviation from a reference. Forexample, the predicted proprotor nacelle trim value may include, orcorrespond to, the predicted propulsor trim value 444 of FIG. 4 or thepredicted proprotor nacelle angle trim value 662 of FIG. 6. The targetstate may include, or correspond to, a target horizontal state (e.g., anairspeed hold state or an acceleration hold state). The aircraftvelocity may include, or correspond to, the aircraft velocity 422 ofFIG. 4 or the vertical velocity 534 of FIG. 5, and the pitch attitudedeviation from the reference may include, or correspond to, the aircraftpitch attitude command 550 of FIG. 5 or the gravity compensatedlongitudinal acceleration signal 622 of FIG. 6.

The method 1600 includes, at 1604, adjusting a nacelle angle of aproprotor of the aircraft based on the predicted proprotor nacelle trimvalue. For example, the propulsor actuator 116 adjusts a nacelle pitchangle (in the pitch axis) of a proprotor the aircraft 100 of FIGS.3A-3D, as described with reference to FIGS. 1, 3A, 4-6, and 13.

In some implementations, the predicted proprotor nacelle trim valueindicates an estimated value of a pitch angle of a nacelle of theproprotor that produces a magnitude of thrust for an airspeed hold stateor an acceleration hold state of the aircraft. In some implementations,the proprotor nacelle command is output to an actuator and is configuredto cause the actuator to adjust a magnitude of proprotor thrustgenerated by adjusting a direction of the thrust in a pitch axis. In aparticular implementation, the proprotor nacelle command is configuredto cause the aircraft to operate in an airspeed hold state or anacceleration hold state of the aircraft.

In some implementations, the method 1600 further includes generating thepitch attitude deviation from the reference based on an input pitchcommand and a selected pitch trim value. For example, the propulsor trimprediction circuitry 404 generates the aircraft pitch attitude command550 of FIG. 5 and the gravity compensated longitudinal accelerationsignal 622 of FIG. 6 based on the pitch attitude input signal 548 andthe selected pitch attitude trim value 546.

In some implementations, the method 1600 further includes generating apredicted pitch attitude trim value based on an aircraft velocity and apitch attitude of the aircraft. The method 1600 includes adjusting anaircraft pitch attitude command based on a pitch attitude input signaland based on the predicted pitch attitude trim value or a commandedpitch attitude trim value. For example, the pitch trim predictioncircuitry 402 generates the predicted pitch attitude trim value 442based on the aircraft velocity 422 and the measured pitch attitude 424and the pitch command model 506 adjusts the aircraft pitch attitudecommand 550 (e.g., a previous value of the aircraft pitch attitudecommand 550 provided as the outer loop feedback signal 948) based onpitch attitude input signal 548 and based on the predicted pitchattitude trim value 442 or the commanded pitch attitude trim value 544,as described with reference to FIG. 9.

In some implementations, the method 1600 further includes operating in ahigh speed mode. In a particular implementation, operating in the highspeed mode includes generating the aircraft pitch attitude command basedon the predicted pitch attitude trim value and generating the predictedproprotor nacelle angle trim value 662 based on the pitch attitude trimvalue 442. For example, the first switch 528 provides the pitch attitudetrim value 442 to the pitch command model 506 and to the propulsor trimprediction circuitry 404 based on the regime signal 540 indicating ahigh speed mode. The pitch command model 506 generates the aircraftpitch attitude command 550 based on the predicted pitch attitude trimvalue 442 and the propulsor trim prediction circuitry 404 generates thepredicted proprotor nacelle angle trim value 662 based on the predictedpitch attitude trim value 442 (or the gravity compensated longitudinalacceleration signal 622, which is generated based on the predicted pitchattitude trim value 442).

Additionally, operating in the high speed mode may further includegenerating a proprotor nacelle command based on a speed select modeacceleration command and the predicted proprotor nacelle trim value. Forexample, the propulsor limiting circuitry 518 generates the propulsorcommand 452 or the nacelle angle command 652 based on the delta nacelleangle command 660 (which is generated based on the speed select modeacceleration command 552) and the predicted proprotor nacelle angle trimvalue 662, as described with reference to FIGS. 4-6. As another example,the combiner 516 generates the combined propulsor command 564 based onthe speed select mode acceleration command 552 and the predictedproprotor nacelle angle trim value 662.

In a particular implementation, generating the speed select modeacceleration command includes generating a pitch independent speedcommand based on a pitch independent acceleration command. The pitchindependent acceleration command is generated based on one or more pilotinputs. Generating the speed select mode acceleration command alsoincludes generating a speed error signal based on subtracting theaircraft velocity from the pitch independent speed command. Generatingthe speed select mode acceleration command further includes applying again to the speed error signal to generate the speed select modeacceleration command. For example, the speed select circuitry 508generates the pitch independent acceleration command 1032 by performinga table lookup using the v-dot command signal 542 (e.g., thumbwheelinceptor input). The speed select circuitry 508 generates the pitchindependent acceleration command error signal 1034 based on theindependent acceleration command 1032 and integrates the pitchindependent acceleration command error signal 1034 to generate the pitchindependent speed command 1038. The speed select circuitry 508 generatesthe speed error signal 1040 based on subtracting the aircraft velocity422 from the pitch independent speed command 1038 and applies thevelocity error gain F(verr) to the speed error signal 1040 to generatethe speed select mode acceleration command 552, as described withreference to FIG. 10.

In some implementations, the method 1600 further includes operating in alow speed mode. In a particular implementation, operating in the lowspeed mode includes generating the aircraft pitch attitude command basedon the commanded pitch attitude trim value and generating the predictedproprotor nacelle trim value based on the commanded pitch attitude trimvalue. For example, the first switch 528 provides the commanded pitchattitude trim value 544 to the pitch command model 506 and to thepropulsor trim prediction circuitry 404 based on the regime signal 540indicating a low speed mode. The pitch command model 506 generates theaircraft pitch attitude command 550 based on the commanded pitchattitude trim value 544, and the propulsor trim prediction circuitry 404generates the predicted proprotor nacelle angle trim value 662 based onthe commanded pitch attitude trim value 544 (or the gravity compensatedlongitudinal acceleration signal 622, which is generated based on thecommanded pitch attitude trim value 544).

Additionally, operating in the low speed mode may include generating anacceleration command mode acceleration command based on the commandedpitch attitude trim value and generating a proprotor nacelle commandbased on the acceleration command mode acceleration command and thepredicted proprotor nacelle trim value. For example, the first switch528 provides the commanded pitch attitude trim value 544 accelerationcommand circuitry 510, and the acceleration command circuitry 510generates the acceleration command mode acceleration command 554 basedon the commanded pitch attitude trim value 544, which is used togenerate the nacelle angle command 652, as described with reference toFIGS. 6 and 11.

In a particular implementation, generating the acceleration command modeacceleration command includes generating a pitch independentacceleration command by performing a table lookup based on a pilotinput, where the pitch independent acceleration command is generatedbased on the one or more pilot inputs. Generating the accelerationcommand mode acceleration command also includes generating a gravityamplified longitudinal acceleration command by performing a table lookupbased on a gravity compensated longitudinal acceleration. Generating theacceleration command mode acceleration command further includesgenerating the acceleration command mode acceleration command based onthe pitch independent acceleration command and the gravity amplifiedlongitudinal acceleration command. For example, the acceleration commandcircuitry 510 generates the pitch independent acceleration command 1032by performing a table lookup using the v-dot command signal 542 (e.g.,thumbwheel inceptor input). The acceleration command circuitry 510generates the acceleration command mode acceleration command 554 basedon combining the pitch independent acceleration command 1032 and thegravity compensated command 1142 generated by the gravity amplifier1124.

FIG. 17 illustrates a particular example of a method 1700 forcontrolling an aircraft, such as the aircraft 100 of FIGS. 1, 2A-2D, and3A-3D. The method 1700 may be performed by a control circuitry orcomponents thereof, such as the control circuitry 130 of FIGS. 1 and4-14.

The method 1700 includes, at 1702, generating a predicted pitch attitudetrim value for a target state of an aircraft based on an aircraftvelocity and a pitch attitude of the aircraft. For example, thepredicted pitch attitude trim value may include or correspond to thepredicted pitch attitude trim value 442 of FIG. 4. The aircraft velocitymay include or correspond to the aircraft velocity 422 of FIG. 4, andthe pitch attitude may include or correspond to the measured pitchattitude 424 of FIG. 4. To illustrate, the pitch trim predictioncircuitry 402 generates the predicted pitch attitude trim value 442based on the aircraft velocity 422 and the measured pitch attitude 424for a target vertical state (e.g., an altitude hold state or a verticalspeed hold state), as described with reference to FIGS. 4-8.

The method 1700 includes, at 1704, adjusting an aircraft pitch attitudecommand based on the predicated pitch attitude trim value and a pilotinput signal from a pitch control inceptor. For example, the aircraftpitch attitude command may include or correspond to the pitch attitudecommand 454 of FIG. 4, the aircraft pitch attitude command 550, or thecontrol surface pitch command 558 of FIG. 5. To illustrate, the controlsurface actuator 410 of FIG. 4 adjusts the control surfaces 128 based onthe pitch attitude command 454 and the second pilot input 414 (e.g., thepitch maneuver input) or a pitch control inceptor input used to generatethe pitch attitude input signal 548 from the pitch control inceptor 526,as described with reference to FIGS. 4 and 5. Alternatively, the pitchactuators 520 of FIG. 5 adjust the control surfaces 128 based on thecontrol surface pitch command 558, as described with reference to FIGS.5-8. In a particular implementation, the aircraft pitch attitude commandis output by output circuitry that includes or corresponds to the pitchcommand model 506, the pitch controller 514, or a combination thereof.

In some implementations, generating the aircraft pitch attitude commandincludes generating an aircraft pitch attitude command error signalbased on subtracting commanded aircraft pitch attitude feedback from aselected aircraft trim pitch attitude value. Generating the aircraftpitch attitude command also includes generating a limited aircraft pitchrate command value (e.g., the limited aircraft pitch rate command 936 ofFIG. 9) based on the aircraft pitch attitude command error signal.Generating the aircraft pitch attitude command further includesgenerating an aircraft pitch rate command error signal based on thelimited aircraft pitch rate command value, a pitch attitude inputsignal, and pitch trim feedback, where the aircraft pitch attitudecommand is generated based on the aircraft pitch rate command errorsignal. For example, the pitch command model 506 generates the aircraftpitch attitude command 550 as described with reference to FIG. 9.

In some implementations, the method 1700 further includes outputting thepredicted pitch attitude trim value to a propulsor trim predictioncircuitry, the propulsor trim prediction circuitry configured togenerate a predicted propulsor collective blade pitch trim value basedon the predicted pitch attitude trim value. For example, the pitch trimprediction circuitry 402 outputs the predicted pitch attitude trim value442 to the propulsor trim prediction circuitry 404 which generates thepredicted propulsor collective blade pitch trim value 562 based on thepredicted pitch attitude trim value 442.

In some implementations, the method 1700 further includes outputting thepitch attitude trim value to an acceleration command circuitry, theacceleration command circuitry configured to generate an accelerationcommand based on the predicted pitch attitude trim value. For example,the pitch trim prediction circuitry 402 outputs the pitch attitude trimvalue 442 to the acceleration command circuitry 510, which generates theacceleration command mode acceleration command 554 based on thepredicted pitch attitude trim value 442.

FIG. 18 illustrates a particular example of a method 1800 forcontrolling an aircraft, such as the aircraft 100 of FIG. 1. The method1800 may be performed by a control circuitry or components thereof, suchas the control circuitry 130 of FIGS. 1 and 4-14, the gravitycompensator 612-618 of FIG. 6, or the gravity amplifier 1124 of FIG. 11or a combination thereof.

The method 1800 includes, at 1802, generating a longitudinalacceleration command as a function of pitch attitude deviation from areference pitch attitude. For example, the longitudinal accelerationcommand may include or correspond to the gravity compensated command1142 of FIG. 11. The pitch attitude deviation may include or correspondto the aircraft pitch attitude command 550 of FIG. 5 or the gravitycompensated longitudinal acceleration signal 622 of FIG. 6. Toillustrate, the gravity amplifier 1124 generates the gravity compensatedcommand 1142 by performing a table lookup using the gravity compensatedlongitudinal acceleration signal 622.

The method 1800 includes, at 1804, adjusting a longitudinal thrusteffector of the aircraft based on the longitudinal acceleration command.For example, the propulsor actuator 116 adjusts a collective blade pitchangle of a propulsor (e.g., a propeller) of the aircraft 100 of FIGS.2A-2D or a nacelle angle of a proprotor of the aircraft 100 of FIGS.3A-3D based on the gravity compensated command 1142, as described withreference to FIGS. 1, 2A, 3A, 4, 5, and 13. Alternatively, the propulsoractuator 116 adjusts a nacelle angular rate, a nozzle size, a nozzledirection, a fuel flow rate, a bypass ratio, thrust bleeding, or thrustvectoring, etc. of one or more of the propulsors 112, 114 of theaircraft 100 based on the gravity compensated command 1142.

FIG. 19 illustrates a particular example of a method 1900 forcontrolling an aircraft, such as the aircraft 100 of FIGS. 1, 2A-2D, and3A-3D. The method 1900 may be performed by a control circuitry orcomponents thereof, such as the control circuitry 130 of FIGS. 1 and4-14.

The method 1900 includes, at 1902, receiving a vertical velocity of theaircraft. For example, the vertical velocity may include or correspondto the vertical velocity 534 of FIG. 5. The method 1900 includes, at1904, receiving a horizontal velocity of the aircraft. For example, thevertical velocity may include or correspond to the aircraft velocity 422(e.g., the horizontal velocity) of FIG. 4. To illustrate, the pitch trimprediction circuitry 402 receives the vertical velocity 534 and theaircraft velocity 422 (e.g., the horizontal velocity) from the sensors132 of FIG. 1 or from the FCC 126 of FIG. 1.

The method 1900 includes, at 1906, filtering a component of the verticalvelocity of the aircraft to generate a filtered vertical velocity. Forexample, the first low pass filter 704 and the first combiner 706function to high pass filter the inverted vertical velocity 732, asdescribed with reference to FIGS. 7 and 8.

The method 1900 includes, at 1908, filtering a measured pitch attitudeof the aircraft to generate a filtered pitch attitude. For example, thesecond low pass filter 724 low pass filters the measured pitch attitude424 to generate the low pass filtered pitch attitude signal 752 (a lowpass filtered measured pitch attitude signal), as described withreference to FIGS. 7 and 8.

The method 1900 includes, at 1910, generating a predicted pitch attitudetrim value for a target vertical state, the predicted pitch attitudetrim value generated based on the horizontal velocity, the filteredvertical velocity, and the filtered pitch attitude. For example, thefourth combiner 730 generates the predicted pitch attitude trim value442 based on combining at least the filtered aircraft trim pitchdeviation signal 750 (generated based on the high pass filtered verticalvelocity signal 736 and the aircraft velocity 422 (e.g., the horizontalvelocity)) and the low pass filtered pitch attitude signal 752.

The method 1900 includes, at 1912, adjusting a flight control effectorbased on the predicted pitch attitude trim value. For example, thecontrol surface actuator 410 or the pitch actuator 520 adjusts one ormore control surfaces 128 based on the pitch attitude command 454 or thecontrol surface pitch command 558 (e.g., an aircraft pitch controlsurface aircraft pitch attitude command) respectively, which aregenerated based on the predicted pitch attitude trim value 442. In someimplementations, the flight control effector includes or corresponds toa flight control surface (e.g., the control surfaces 128) or a propulsor(e.g., the propulsors 112, 114). In a particular implementation, thetarget vertical state is a vertical speed hold state or an altitude holdstate.

In some implementations, the flight control effector includes orcorresponds to a flight control surface (e.g., the control surface 128)of the aircraft. Additionally or alternatively, the flight controlsurface includes or corresponds to an elevator, a flaps, a slat, anaileron (e.g., a flaperon), a spoiler, a tab, or another pitch attitudecontrol surface. In such implementations, adjusting the flight controlsurface based on the predicted pitch attitude trim value includesgenerating a pitch attitude command for the target vertical state basedon the predicted pitch attitude trim value and adjusting the flightcontrol surface of the aircraft based on the pitch attitude command. Forexample, the processing circuitry 408 generates the pitch attitudecommand 454 for the target vertical state based on the predicted pitchattitude trim value 442 and the control surface actuator 410 adjusts thecontrol surface 128 of the aircraft based on the pitch attitude command454.

In some implementations, the method 1900 further includes generating alongitudinal thrust effector command for a target horizontal state basedon the predicted pitch attitude trim value and adjusting a longitudinalthrust effector of the aircraft based on the longitudinal thrusteffector command. For example, the processing circuitry 406 generatesthe propulsor command 452 for a target horizontal state based on thepredicted pitch attitude trim value 442 and the propulsor actuator 116adjusts one or more of the propulsors 112, 114 of the aircraft 100 basedon the propulsor command 452.

In some implementations, the flight control effector is a longitudinalthrust effector of the aircraft and the method 1900 further includesgenerating a predicted longitudinal thrust effector trim value for atarget horizontal state based on the predicted pitch attitude trimvalue, generating a longitudinal thrust effector command for the targethorizontal state based on the predicted longitudinal thrust effectortrim value, and adjusting the longitudinal thrust effector of theaircraft based on the longitudinal thrust effector command. For example,the propulsor trim prediction circuitry 404 generates the predictedpropulsor trim value (e.g., one of the trim values 562, 662) for atarget horizontal state based on the predicted pitch attitude trim value442, the propulsor limiting circuitry 518 generates the propulsorcommand 452 for the target horizontal state based on the combinedpropulsor command 564 (which is generated based on the predictedpropulsor trim value (e.g., one of the trim values 562, 662)), and thepropulsor actuator 116 adjusts one or more of the propulsors 112, 114 ofthe aircraft 100 based on the propulsor command 452. In someimplementation, the longitudinal thrust effector (e.g., a propulsor)includes or corresponds a propeller, a proprotor, a ducted fan, acontra-rotating fan, a turbojet engine, a turbofan engine, or anotherlongitudinal thrust effector.

In some implementations, the method 1900 further includes generating asecond predicted pitch attitude trim value for a second target verticalstate and generating a commanded pitch attitude trim value based on apilot input. For example, the pitch trim prediction circuitry 402generates a second predicted pitch attitude trim value 442 for a secondtarget vertical state and the integrator 504 generates the commandedpitch attitude trim value 544 based on a pilot input (e.g., a pitch triminceptor input) to the pitch trim inceptor 524. The second predictedpitch attitude trim value 442 is generated at a second time that isafter a first time when the predicted pitch attitude trim value 442 wasgenerated. In a particular implementation, the method 1900 also includesselecting, based on a regime control signal, the commanded pitchattitude trim value as a selected pitch attitude trim value andadjusting the flight control effector based on the selected pitchattitude trim value and independent of the second predicted pitchattitude trim value. For example, the first switch 528 selects, based onthe regime signal 540, the commanded pitch attitude trim value 544 andoutputs the commanded pitch attitude trim value 544 as the selectedpitch attitude trim value 546. A particular pitch actuator 520 adjusts aparticular control surface 128 based on the selected pitch attitude trimvalue 546 and independent of the second predicted pitch attitude trimvalue 442. Additionally or alternatively, a particular propulsoractuator 116 adjusts a particular propulsor based on the selected pitchattitude trim value 546 and independent of the second predicted pitchattitude trim value 442.

In some implementations, the method 1900 further includes generating avertical flight path angle signal based on a selected airspeed and thefiltered vertical velocity. For example, the first multiplier 708generates the vertical flight path angle signal 744 based on the inverseselected speed 748 and the low pass filtered vertical velocity signal734. As another example, the divider 802 generates the vertical flightpath angle signal 744 based on the selected speed 746 and the low passfiltered vertical velocity signal 734.

In some implementations, the method 1900 further includes filtering thehigh pass filtered vertical velocity to generate a filtered verticalacceleration, dividing the filtered vertical acceleration by an invertedselected speed to generate a time rate of change of flight path anglesignal, multiplying the time rate of change of flight path angle signalby a vertical damping derivative to generate a damped signal, where thepredicted pitch attitude trim value is generated further based on thedamped signal. For example, the circuitry 710 applies a gain of theinverse of the time constant to the high pass filtered vertical velocitysignal 736 to generate the filtered vertical acceleration 738. Themultiple 712 multiplies the filtered vertical acceleration 738 by theinverse selected speed 748 to generate the time rate of change of flightpath angle signal 740. The vertical damping derivative (Zw) circuitry714 divides the time rate of change of flight path angle signal 740 by avertical damping derivative (Zw) to generate the first signal 742, whichis used to generate the predicted pitch attitude trim value 442.

In some implementations, the method 1900 further includes generating afiltered aircraft trim pitch deviation signal based on combining avertical flight path angle and the first signal, where generating thepredicted pitch attitude trim value for the target vertical stateincludes combining at least the filtered aircraft trim pitch deviationsignal and the filtered pitch attitude. For example, the second combiner716 generates the filtered aircraft trim pitch deviation signal 750based on combining the vertical flight path angle signal 744 and thefirst signal 742. The fourth combiner 730 generates the predicted pitchattitude trim value 442 for the target vertical state by combining atleast the filtered aircraft trim pitch deviation signal 750 and the lowpass filtered pitch attitude signal 752.

In some implementations, the method 1900 further includes filtering thehigh pass vertical velocity to generate a filtered vertical accelerationand dividing the filtered vertical acceleration by a normalizeddimensional aircraft vertical force per angle of attack derivative togenerate a normalized signal, where the predicted pitch attitude trimvalue is generated further based on the normalized signal. For example,the gain circuitry 710 multiplies the high pass filtered verticalvelocity signal 736 by a gain of the inverse of the time constant togenerate the filtered vertical acceleration 738 and the normalizeddimensional aircraft vertical force per angle of attack derivative (Zα)circuitry 804 divides the filtered vertical acceleration 738 by thenormalized dimensional aircraft vertical force per angle of attackderivative (Zα) to generate the normalized signal 812.

In some implementations, the method 1900 further includes generating afiltered aircraft trim pitch deviation signal based on combining avertical flight path angle and the normalized signal, where generatingthe predicted pitch attitude trim value for the target vertical stateincludes combining at least the filtered aircraft trim pitch deviationsignal and the filtered pitch attitude. For example, the second combiner716 generates the filtered aircraft trim pitch deviation signal 750based on combining the vertical flight path angle signal 744 and thenormalized signal 812. The fourth combiner 730 generates the predictedpitch attitude trim value 442 for the target vertical state based oncombining at least the filtered aircraft trim pitch deviation signal 750and the filtered pitch attitude signal 752.

In some implementations, generating the aircraft pitch attitude commandincludes generating an aircraft pitch attitude command error signalbased on subtracting commanded aircraft pitch attitude feedback from aselected aircraft trim pitch attitude value, generating a limitedaircraft pitch rate command value based on the aircraft pitch attitudecommand error signal and generating an aircraft pitch rate command errorsignal based on the limited aircraft pitch rate command value, a pitchattitude input signal, and pitch trim feedback, where the aircraft pitchattitude command is generated based on the aircraft pitch rate commanderror signal. For example, the first combiner 914 generates the aircraftpitch attitude command error signal 932 based on subtracting the outerloop feedback signal 948 from the selected pitch attitude trim value546. The first authority limiter 918 generates the limited aircraftpitch rate command 936 based on the aircraft pitch attitude commanderror signal 932. The second combiner 920 generates the aircraft pitchrate command error signal 940 based on the limited aircraft pitch ratecommand 936, the pitch attitude input signal 548, and the inner loopfeedback signal 938 (pitch trim feedback). The pitch command modelgenerates the aircraft pitch attitude command 550 based on the aircraftpitch rate command error signal 940, as described with reference to FIG.9. The aircraft pitch attitude command 550 indicates a pitch attitudedeviation from a reference pitch attitude (e.g., a deviation from thepitch attitude input signal 548 by the predicted pitch attitude trimvalue 442 or the commanded pitch attitude trim value 544 of the selectedpitch attitude trim value 546).

In some implementations, generating the aircraft pitch attitude commandincludes generating an aircraft pitch attitude command error signalbased on subtracting commanded aircraft pitch attitude feedback from aselected aircraft trim pitch attitude value. Generating the aircraftpitch attitude command includes generating an aircraft pitch trim ratefeedback signal by applying a gain to the aircraft pitch attitudecommand error signal. Generating the aircraft pitch attitude commandalso includes generating a limited aircraft pitch rate command valuebased on limiting the aircraft pitch trim rate feedback signal.Generating the aircraft pitch attitude command includes generating anaircraft pitch rate command error signal based on the limited aircraftpitch rate command value, a pitch attitude input signal, and pitch trimfeedback. Generating the aircraft pitch attitude command also includesamplifying the aircraft pitch rate command error signal to generate anamplified aircraft pitch rate command error signal. Generating theaircraft pitch attitude command includes generating a pitch angularacceleration command based on limiting the amplified aircraft pitch ratecommand error signal. Generating the aircraft pitch attitude commandalso includes integrating the pitch angular acceleration command togenerate a pitch rate command. Generating the aircraft pitch attitudecommand includes applying a coordination transformation to the pitchrate command to generate a pitch attitude rate command. Generating theaircraft pitch attitude command further includes integrating the pitchattitude rate command to generate the aircraft pitch attitude command.For example, the pitch command model 506 generates the aircraft pitchattitude command 550 as described with reference to FIG. 9.

FIG. 20 illustrates a particular example of a method 2000 of method forcontrolling an aircraft, such as the aircraft 100 of FIGS. 1, 2A-2D, and3A-3D. The method 2000 may be performed by a control circuitry orcomponents thereof, such as the control circuitry 130 of FIGS. 1 and4-14.

The method 2000 includes, at 2002, filtering a gravity compensatedlongitudinal acceleration of the aircraft to generate a filtered gravitycompensated longitudinal acceleration. For example, the gravitycompensated longitudinal acceleration may include, or correspond to, thegravity compensated longitudinal acceleration signal 622 of FIG. 6. Thefiltered gravity compensated longitudinal acceleration may include, orcorrespond to, the high pass filtered gravity compensated longitudinalacceleration signal 1344 of FIG. 13. To illustrate, the first low passfilter 1312 and the first combiner 1322 function to high pass filter thegravity compensated longitudinal acceleration signal 622 to generate thehigh pass filtered gravity compensated longitudinal acceleration signal1344, as described with reference to FIG. 13.

The method 2000 includes, at 2004, filtering a speed of the aircraft togenerate a filtered speed of the aircraft. For example, the speed mayinclude, or correspond to, the aircraft velocity 422 of FIG. 4, such asa horizontal velocity. To illustrate, the second low pass filter 1314and the second combiner 1324 function to high pass filter the aircraftvelocity 422 to generate the high pass filtered aircraft velocity 1348,as described with reference to FIG. 13.

The method 2000 includes, at 2006, generating a filtered longitudinalcontrol effector error based on the filtered gravity compensatedlongitudinal acceleration and the filtered speed. For example, the thirdcombiner 1326 combines the high pass filtered gravity compensatedlongitudinal acceleration signal 1344 and the filtered accelerationfeedback 1350 (which is generated based on the high pass filteredaircraft velocity 1348) to generate the offset value 1352 and thedivider 1332 divides the offset value 1352 by the propulsor sensitivityvalue 1242 to generate the filtered longitudinal control effector error1356, as described with reference to FIG. 13.

The method 2000 includes, at 2008, filtering a longitudinal thrusteffector command value to generate a filtered longitudinal thrusteffector command value. For example, the longitudinal thrust effectorcommand value includes, or corresponds to, the propulsor feedback value432 of FIG. 4, the propulsor collective blade pitch feedback 566 of FIG.5, or the nacelle angle feedback 664 of FIG. 6. To illustrate, the thirdlow pass filter 1316 low pass filters the propulsor collective bladepitch feedback 566 to generate the low pass filtered propulsorcollective blade pitch feedback 1358, as described with reference toFIG. 13. In other implementations, the filtered longitudinal thrusteffector command value is a low pass filtered longitudinal thrusteffector command value generated based on a measured value determinedbased on sensor data, as opposed to feedback value, as explained withreference to FIG. 4.

The method 2000 includes, at 2010, generating a predicted longitudinalthrust effector trim value for a target horizontal state, the predictedlongitudinal thrust effector trim value generated based on the filteredlongitudinal control effector error and the filtered longitudinal thrusteffector command value. For example, the predicted longitudinal thrusteffector trim value includes or corresponds to the predicted propulsortrim value 444 of FIG. 4, the predicted propulsor collective blade pitchtrim value 562 of FIG. 5, or the predicted proprotor nacelle angle trimvalue 662 of FIG. 6. To illustrate, the fourth combiner 1328 generatesthe 562 based on combining the filtered longitudinal control effectorerror 1356 (generated based on the offset value 1352) and the low passfiltered propulsor collective blade pitch feedback 1358.

The method 2000 includes, at 2012, adjusting a longitudinal thrusteffector of the aircraft based on the predicted longitudinal thrusteffector trim value. For example, the propulsor actuator 116 adjusts oneor more components of the propulsors 112, 114 based on the propulsorcommand 452 (which is generated based on the predicted propulsor trimvalue 444), as described with reference to FIG. 4.

In some implementations, the longitudinal thrust effector of theaircraft includes or corresponds to a propeller, a proprotor, a rotor, aducted fan, a contra-rotating propeller, a turbojet engine, a turbofanengine, or a rocket. In a particular implementation, the targethorizontal state includes or corresponds to an airspeed hold state or anacceleration hold state.

In some implementations, the longitudinal thrust effector is adjusted byan actuator. For example, the propulsors 112, 114 are adjusted by thepropulsor actuator 116, as described with reference to FIG. 1. In aparticular implementation, the actuator includes, or corresponds to, apropulsor collective actuator, a propulsor cyclic actuator, a nacelleactuator, a propulsor nozzle actuator, or a fuel flow rate actuator.Additionally or alternatively, adjusting the longitudinal thrusteffector causes the aircraft to operate in the airspeed hold state orthe acceleration hold state.

In a particular implementation, generating the gravity compensatedlongitudinal acceleration includes generating a pitch attitude deviationfrom a reference by subtracting a sine value of a selected trim pitchattitude value and a sine value of a commanded pitch attitude andgenerating the gravity compensated longitudinal acceleration bymultiplying the pitch attitude deviation from the reference by anacceleration due to gravity constant. For example, the combiner 616generates the longitudinal acceleration signal 1138 by subtracting thesine value of the selected pitch attitude trim value 1134 and the sineof the aircraft pitch attitude command 1136. The gravity multiplier 618generates the gravity compensated longitudinal acceleration signal 622by multiplying the longitudinal acceleration signal 1138 by theacceleration due to gravity constant (e.g., approximately 9.8 meters persecond squared).

In some implementations, the selected pitch attitude trim value 546 isgenerated based on a predicted pitch attitude trim value 442 for atarget vertical state, such as when the control circuitry 130 isoperating in a high speed mode and the regime signal 540 indicates thehigh speed mode, as described with reference to FIG. 5.

In some implementations, the longitudinal thrust effector command valueis a measured value determined based on sensor data or a feedback valueincluding a prior longitudinal thrust effector command value. Forexample, the propulsor feedback value 432 is either generated based onsensor data from the sensors 132 of FIG. 1 or is generated based on aprevious value (e.g., a feedback value) of the propulsor command 452, asdescribed with reference to FIG. 4.

In some implementations, the method 2000 further includes generating anoffset value based on subtracting the filtered acceleration feedbackfrom the filtered gravity compensated longitudinal acceleration, wheregenerating the filtered longitudinal control effector error includesdividing the offset value by a propeller sensitivity value, and wheregenerating the predicted longitudinal thrust effector trim valueincludes combining the filtered longitudinal control effector error andthe filtered longitudinal thrust effector command value. For example,the third combiner 1326 generates the offset value 1352 by subtractingthe filtered acceleration feedback 1350 from the filtered gravitycompensated longitudinal acceleration signal 1344 and divider 1332divides the offset value by the propulsor sensitivity value 1242 togenerate the filtered longitudinal control effector error 1356. Thefourth combiner 1328 generates the predicted propulsor collective bladepitch trim value 562 by combining the filtered longitudinal controleffector error 1356 and the low pass filtered propulsor collective bladepitch feedback 1358.

In some implementations, the method 2000 further includes generating thepropeller sensitive value by performing a table lookup to a propellersensitivity schedule based on the speed of the aircraft or a propellerrotational speed. For example, the propulsor sensitivity scheduler 1222generates the propulsor sensitivity value 1242 by performing a tablelookup to the propeller sensitivity schedule based on the aircraftvelocity 422.

In some implementations, the method 2000 further includes generating alongitudinal thrust effector command for the target horizontal state. Ina particular implementation, generating the longitudinal thrust effectorcommand for the target horizontal state includes generating thepredicted longitudinal thrust effector trim value based on combining afiltered longitudinal control effector error and the filteredlongitudinal thrust effector command value, the filtered longitudinalcontrol effector error generated based on the offset value, combining adelta propulsor command with the predicted longitudinal thrust effectortrim value to generate a combined propulsor command, and limiting thecombined propulsor command to generate the longitudinal thrust effectorcommand. For example, the fourth combiner 1328 generates the predictedpropulsor collective blade pitch trim value 562 based on combining thefiltered longitudinal control effector error 1356 and the low passfiltered propulsor collective blade pitch feedback 1358, the combiner516 combines the delta propulsor collective blade pitch command 560 withthe predicted propulsor collective blade pitch trim value 562 togenerate the combined propulsor command 564, and the propulsor limitingcircuitry 518 limits the combined propulsor command 564 to generate thepropulsor command 452.

In some implementations, the delta propulsor command is generated basedon the speed select mode acceleration command. For example, theacceleration controller 512 receives the speed select mode accelerationcommand from the second switch 530 based on the regime signal 540indicating the high speed regime or mode. In such implementations,generating the speed select mode acceleration command 552 includesperforming a table lookup based on a pilot input to generate a pitchindependent acceleration command and generating a pitch independentspeed command based on the pitch independent acceleration command. Forexample, the speed select circuitry 508 performs a table lookup based onthe v-dot command signal 542 (which is generated based on a pilot inputto the thrust inceptor 522) to generate the pitch independentacceleration command 1032 and the integrator 1022 generates the pitchindependent speed command 1038 based on the pitch independentacceleration command 1032 by integrating the pitch independentacceleration command error signal 1034. Generating the speed select modeacceleration command 552 further includes generating a speed errorsignal based on subtracting the speed of the aircraft from the pitchindependent speed command, applying a gain to the speed error signal togenerate a speed feedback acceleration command, and outputting the speedselect mode acceleration command to a switch. For example, the secondcombiner 1024 generates the speed error signal 1040 based on subtractingthe aircraft velocity 422 from the pitch independent speed command 1038,and the second amplifier 1026 applies the velocity error gain (F(verr))to the speed error signal 1040 to generate the speed select modeacceleration command 552 and outputs the speed select mode accelerationcommand 552 to the second switch 530.

In some implementations, generating the speed select mode accelerationcommand further includes generating a washout signal based on the speederror signal and generating a pitch independent acceleration commanderror signal based on subtracting the washout signal from the pitchindependent acceleration command. For example, the third amplifier 1030(e.g., a washout amplifier) receives the limited speed error feedbacksignal 1044 and applies the washout gain Kwo to the limited speed errorfeedback signal 1044 to generate the washout signal 1046 and the firstcombiner 1016 subtracts the washout signal 1046 from the pitchindependent acceleration command 1032 to generate the pitch independentacceleration command error signal 1034. In a particular implementation,generating the pitch independent speed command 1038 includes integratingthe pitch independent acceleration command error signal 1034 to generatethe pitch independent speed command 1038.

In some implementations, the delta propulsor command is generated basedon the acceleration command mode acceleration command, and the method2000 further includes generating the acceleration command modeacceleration command based on a pitch independent acceleration commandand a gravity amplified longitudinal acceleration command. For example,the acceleration command circuitry 510 generates the accelerationcommand mode acceleration command 554 based on the pitch independentacceleration command 1032 and a gravity compensated command 1142, asdescribed with reference to FIG. 11. In a particular implementation, thepitch independent acceleration command 1032 is generated based on apilot input. For example, the thrust inceptor 522 receives a pilot inputand generates the v-dot command signal 542 based on the pilot input.

In some implementations, the method 2000 further includes providing anacceleration command to a combiner as an acceleration feedforwardcommand. For example, the acceleration controller 512 provides theselected acceleration command 556 to the second combiner 1218 as thefeedforward acceleration command 1232. The acceleration commandcorresponds to the speed select mode acceleration command 552 or theacceleration command mode acceleration command 554. The method 2000includes generating an acceleration error signal based on theacceleration command and a measured acceleration of the aircraft. Forexample, the first combiner 1212 generates the acceleration error signal1234 based on combining the selected acceleration command 556 and themeasured acceleration of the aircraft, i.e., the input v-dot 1230. Themethod 2000 includes applying gain to the acceleration error signal togenerate an acceleration feedback command. For example, the firstamplifier 1214 applies acceleration feedback gain (Kvdfb) to theacceleration error signal 1234 to generate the acceleration feedbackcommand 1236. The method 2000 includes authority limiting theacceleration feedback command to generate a limited accelerationfeedback command. For example, the authority limiter 1216 limits theacceleration feedback command 1236 to generate the limited accelerationfeedback command 1238 based on one or more authority limit thresholds.The method 2000 includes generating, by the combiner, a combinedacceleration command based on the limited acceleration feedback commandand the feedforward acceleration command. For example, the combiner 1218generates the combined acceleration command 1240 based on combining thelimited acceleration feedback command 1238 and the feedforwardacceleration command 1232. The method 2000 includes generating the deltapropulsor command based on dividing the combined acceleration command bya propulsor thrust sensitivity value, where the propulsor command isgenerated further based on the delta propulsor command. For example, thedivider 1220 generates the delta propulsor collective blade pitchcommand 560 based on dividing the combined acceleration command 1240 bythe propulsor sensitivity value 1242.

The methods 1500-2000 of FIGS. 15-20 may be initiated or controlled byan application-specific integrated circuit (ASIC), a processing unit,such as a central processing unit (CPU), a controller, another hardwaredevice, a firmware device, a field-programmable gate array (FPGA)device, or any combination thereof. As an example, the method 1500 ofFIG. 15 can be initiated or controlled by one or more processors, suchas one or more processors included in a control system. In someimplementations, a portion of one of the methods FIGS. 15-20 may becombined with a second portion of one of the methods of FIGS. 15-20.Additionally, one or more operations described with reference to theFIGS. 15-20 may be optional, may be performed at least partiallyconcurrently, and/or may be performed in a different order than shown ordescribed.

Referring to FIGS. 21 and 22, examples of the disclosure are describedin the context of a vehicle manufacturing and service method 2100 asillustrated by the flow chart of FIG. 21 and a vehicle 2202 asillustrated by the block diagram 2200 of FIG. 22. A vehicle produced bythe vehicle manufacturing and service method 2100 of FIG. 21, such asthe vehicle 2202 of FIG. 22, may include an aircraft, an airship, arocket, a satellite, a submarine, or another vehicle, as illustrative,non-limiting examples. The vehicle 2202 may be manned or unmanned (e.g.,a drone or an unmanned aerial vehicle (UAV)).

Referring to FIG. 21, a flowchart of an illustrative example of a methodof control circuitry manufacturing and service is shown and designated2100. During pre-production, the exemplary method 2100 includes, at2102, specification and design of a vehicle, such as a vehicle 2202described with reference to FIG. 22. During the specification and designof the vehicle 2202, the method 2100 may include specifying a design ofa control circuitry, such as the control circuitry 130 of FIG. 1. At2104, the method 2100 includes material procurement. For example, themethod 2100 may include procuring materials for the control circuitry ofthe vehicle 2202.

During production, the method 2100 includes, at 2106, component andsubassembly manufacturing and, at 2108, system integration of thevehicle 2202. The method 2100 may include component and subassemblymanufacturing (e.g., manufacturing and or programming the controlcircuitry 130 of FIG. 1) of the vehicle 2202 and system integration(e.g., coupling the control circuitry 130 of FIG. 1 to one or morecomponents of the vehicle 2202). At 2110, the method 2100 includescertification and delivery of the vehicle 2202 and, at 2112, placing thevehicle 2202 in service. Certification and delivery may includecertifying the control circuitry 130 of FIG. 1 by inspection ornon-destructive testing. While in service by a customer, the vehicle2202 may be scheduled for routine maintenance and service (which mayalso include modification, reconfiguration, refurbishment, and so on).At 2114, the method 2100 includes performing maintenance and service onthe vehicle 2202. The method 2100 may include performing maintenance andservice of the control circuitry 130. For example, maintenance andservice of the communications system may include replacing the controlcircuitry 130 or updating the control circuitry 130.

Each of the processes of the method 2100 may be performed or carried outby a system integrator, a third party, and/or an operator (e.g., acustomer). For the purposes of this description, a system integrator mayinclude without limitation any number of vehicle manufacturers andmajor-system subcontractors; a third party may include withoutlimitation any number of vendors, subcontractors, and suppliers; and anoperator may be an airline, leasing company, military entity, serviceorganization, and so on.

Referring to FIG. 22, a block diagram 2200 of an illustrativeimplementation of the vehicle 2202 that includes a control circuitry,such as the control circuitry 130 of FIG. 1. To illustrate, the vehicle2202 may include an aircraft, as an illustrative, non-limiting example.The vehicle 2202 may have been produced by at least a portion of themethod 2100 of FIG. 21. As shown in FIG. 22, the vehicle 2202 (e.g., theaircraft 100 of FIG. 1) includes an airframe 2218, an interior 2222, thecontrol circuitry 130, and a plurality of systems 2220. The plurality ofsystems 2220 may include one or more of a propulsion system 2224, anelectrical system 2226, an environmental system 2228, or a hydraulicsystem 2230. The control circuitry 130 may include the pitch trimprediction circuitry 402, the propulsor trim prediction circuitry 404,or both. The control circuitry 130 (or components thereof) may beconfigured to perform one or more steps of the methods 1500-2000 ofFIGS. 15-20 and/or as described with reference to FIG. 1.

Apparatus and methods included herein may be employed during any one ormore of the stages of the method 2100 of FIG. 21. For example,components or subassemblies corresponding to production process 2108 maybe fabricated or manufactured in a manner similar to components orsubassemblies produced while the vehicle 2202 is in service, at 2112 forexample and without limitation. Also, one or more apparatusimplementations, method implementations, or a combination thereof may beutilized during the production stages (e.g., stages 2102-2110 of themethod 2100), for example, by substantially expediting assembly of orreducing the cost of the vehicle 2202. Similarly, one or more ofapparatus implementations, method implementations, or a combinationthereof, may be utilized while the vehicle 2202 is in service, at 2112for example and without limitation, to maintenance and service, at 2114.

The illustrations of the examples described herein are intended toprovide a general understanding of the structure of the variousimplementations. The illustrations are not intended to serve as acomplete description of all of the elements and features of apparatusand systems that utilize the structures or methods described herein.Many other implementations may be apparent to those of skill in the artupon reviewing the disclosure. Other implementations may be utilized andderived from the disclosure, such that structural and logicalsubstitutions and changes may be made without departing from the scopeof the disclosure. For example, method operations may be performed in adifferent order than shown in the figures or one or more methodoperations may be omitted. Accordingly, the disclosure and the figuresare to be regarded as illustrative rather than restrictive.

Moreover, although specific examples have been illustrated and describedherein, it should be appreciated that any subsequent arrangementdesigned to achieve the same or similar results may be substituted forthe specific implementations shown. This disclosure is intended to coverany and all subsequent adaptations or variations of variousimplementations. Combinations of the above implementations, and otherimplementations not specifically described herein, will be apparent tothose of skill in the art upon reviewing the description.

The Abstract of the Disclosure is submitted with the understanding thatit will not be used to interpret or limit the scope or meaning of theclaims. In addition, in the foregoing Detailed Description, variousfeatures may be grouped together or described in a single implementationfor the purpose of streamlining the disclosure. Examples described aboveillustrate but do not limit the disclosure. It should also be understoodthat numerous modifications and variations are possible in accordancewith the principles of the present disclosure. As the following claimsreflect, the claimed subject matter may be directed to less than all ofthe features of any of the disclosed examples. Accordingly, the scope ofthe disclosure is defined by the following claims and their equivalents.

What is claimed is:
 1. A system for an aircraft, the system comprising:a longitudinal thrust effector; control circuitry coupled to thelongitudinal thrust effector, wherein the control circuitry comprises: afirst filter configured to filter a gravity compensated longitudinalacceleration of the aircraft to generate a filtered gravity compensatedlongitudinal acceleration; a second filter configured to generate afiltered speed of the aircraft based on a speed of the aircraft;intermediary circuitry configured to generate a filtered longitudinalcontrol effector error based on the gravity compensated longitudinalacceleration and the filtered speed; a third filter configured togenerate a filtered longitudinal thrust effector command value based ona longitudinal thrust effector command value; and output circuitryconfigured to generate a predicted longitudinal thrust effector trimvalue for a target horizontal state, the predicted longitudinal thrusteffector trim value generated based on the filtered longitudinal controleffector error and the filtered longitudinal thrust effector commandvalue; and an actuator coupled to the control circuitry and thelongitudinal thrust effector, wherein the actuator is configured toreceive a command based on the predicted longitudinal thrust effectortrim value and adjust the longitudinal thrust effector based on thecommand.
 2. The system of claim 1, wherein the longitudinal thrusteffector comprises a propeller, a proprotor, a rotor, a ducted fan, acontra-rotating propeller, a turbojet engine, a turbofan engine, or arocket, and wherein the target horizontal state includes an airspeedhold state or an acceleration hold state.
 3. The system of claim 1,wherein the actuator comprises a propulsor collective actuator, apropulsor cyclic actuator, a nacelle actuator, a propulsor nozzleactuator, or a fuel flow rate actuator, and wherein adjustment of thelongitudinal thrust effector is configured to cause the aircraft tooperate in an airspeed hold state or an acceleration hold state.
 4. Thesystem of claim 1, wherein the first filter comprises a high pass filterconfigured to generate a high pass filtered gravity compensatedlongitudinal acceleration based on high pass filtering the gravitycompensated longitudinal acceleration.
 5. The system of claim 4, whereinthe first filter includes: a low pass filter configured to low passfilter the gravity compensated longitudinal acceleration to generate alow pass filtered gravity compensated longitudinal acceleration; and acombiner configured to subtract the low pass filtered gravitycompensated longitudinal acceleration from the gravity compensatedlongitudinal acceleration to generate the high pass filtered gravitycompensated longitudinal acceleration.
 6. The system of claim 1, whereinthe speed of the aircraft comprises a horizontal velocity of theaircraft, and wherein the second filter comprises a high pass filterconfigured to generate a high pass filtered speed of the aircraft basedon high pass filtering the speed of the aircraft.
 7. The system of claim6, further comprising a gain circuitry configured to generate a filteredacceleration feedback of the aircraft based on multiplying the high passfiltered speed of the aircraft by a gain value of an inverse of a timeconstant of the control circuitry, wherein the filtered longitudinalcontrol effector error is generated based on the filtered accelerationfeedback.
 8. The system of claim 1, wherein the third filter comprises alow pass filter configured to low pass filter the longitudinal thrusteffector command value to generate a low pass filtered longitudinalthrust effector command value.
 9. A method of controlling an aircraft,the method comprising: filtering a gravity compensated longitudinalacceleration of the aircraft to generate a filtered gravity compensatedlongitudinal acceleration; filtering a speed of the aircraft to generatea filtered speed of the aircraft; generating a filtered longitudinalcontrol effector error based on the gravity compensated longitudinalacceleration and the filtered speed; filtering a longitudinal thrusteffector command value to generate a filtered longitudinal thrusteffector command value; generating a predicted longitudinal thrusteffector trim value for a target horizontal state, the predictedlongitudinal thrust effector trim value generated based on the filteredlongitudinal control effector error and the filtered longitudinal thrusteffector command value; and adjusting a longitudinal thrust effector ofthe aircraft based on the predicted longitudinal thrust effector trimvalue.
 10. The method of claim 9, wherein generating the gravitycompensated longitudinal acceleration comprises: applying a sinefunction to a selected trim pitch attitude value to generate a sine ofthe selected trim pitch attitude value; applying a sine function to apitch attitude command to generate a sine of the pitch attitude command;generating a longitudinal acceleration signal by subtracting the sine ofthe selected trim pitch attitude value from the sine of the pitchattitude command; and generating the gravity compensated longitudinalacceleration by multiplying the longitudinal acceleration signal by anacceleration due to gravity constant.
 11. The method of claim 10,wherein the selected trim pitch attitude value is generated based on apredicted pitch attitude trim value for a target vertical state.
 12. Themethod of claim 9, wherein the longitudinal thrust effector commandvalue comprises a measured value determined based on sensor data or afeedback value comprising a prior longitudinal thrust effector commandvalue.
 13. The method of claim 9, further comprising generating anoffset value based on subtracting a filtered acceleration feedback fromthe filtered gravity compensated longitudinal acceleration, whereingenerating the filtered longitudinal control effector error includesdividing the offset value by a propeller sensitivity value, and whereingenerating the predicted longitudinal thrust effector trim valueincludes combining the filtered longitudinal control effector error andthe filtered longitudinal thrust effector command value.
 14. The methodof claim 13, further comprising generating the propeller sensitivityvalue by performing a table lookup to a propeller sensitivity schedulebased on the speed of the aircraft or a propeller rotational speed. 15.The method of claim 9, further comprising generating a longitudinalthrust effector command for the target horizontal state, whereingenerating the longitudinal thrust effector command for the targethorizontal state includes: generating the predicted longitudinal thrusteffector trim value based on combining the filtered longitudinal controleffector error and the filtered longitudinal thrust effector commandvalue; combining a delta propulsor command with the predictedlongitudinal thrust effector trim value to generate a combined propulsorcommand; and limiting the combined propulsor command to generate thelongitudinal thrust effector command, wherein the longitudinal thrusteffector of the aircraft is adjusted based on the longitudinal thrusteffector command.
 16. The method of claim 15, further comprisinggenerating the delta propulsor command based on a speed select modeacceleration command or an acceleration command mode accelerationcommand.
 17. The method of claim 16, wherein the delta propulsor commandis generated based on the speed select mode acceleration command, andfurther comprising generating the speed select mode accelerationcommand, wherein generating the speed select mode acceleration commandincludes: performing a table lookup based on a pilot input to generate apitch independent acceleration command; generating a pitch independentspeed command based on the pitch independent acceleration command;generating a speed error signal based on subtracting the speed of theaircraft from the pitch independent speed command; applying a gain tothe speed error signal to generate the speed select mode accelerationcommand; and outputting the speed select mode acceleration command to aswitch.
 18. The method of claim 17, wherein generating the speed selectmode acceleration command further includes: generating a washout signalbased on the speed error signal; and generating a pitch independentacceleration command error signal based on subtracting the washoutsignal from the pitch independent acceleration command, whereingenerating the pitch independent speed command includes integrating thepitch independent acceleration command error signal to generate a pitchindependent speed command.
 19. The method of claim 16, wherein the deltapropulsor command is generated based on the acceleration command modeacceleration command, and further comprising generating the accelerationcommand mode acceleration command based on a pitch independentacceleration command and a gravity amplified longitudinal accelerationcommand, wherein the pitch independent acceleration command is generatedbased on a pilot input.
 20. The method of claim 15, further comprising:providing an acceleration command to a combiner as an accelerationfeedforward command, wherein the acceleration command corresponds to aspeed select mode acceleration command or an acceleration command modeacceleration command; generating an acceleration error signal based onthe acceleration command and a measured acceleration of the aircraft;applying gain to the acceleration error signal to generate anacceleration feedback command; authority limiting the accelerationfeedback command to generate a limited acceleration feedback command;generating, by the combiner, a combined acceleration command based onthe limited acceleration feedback command and the feedforwardacceleration command; and generating the delta propulsor command basedon dividing the combined acceleration command by a propulsor thrustsensitivity value.